NASA PDR Technical Report

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NASA PDR Technical Report

  1. 1. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review i i Note to reader: To facilitate the reading of the Preliminary Design review, we have mirrored the Student Launch Project Statement of Work. In the body of the PDR, you will find extensive detail in the design of our SMD payload. The payload’s features are threefold with atmospheric data gathering sensors, a self-leveling camera system, and video camera. One of the two major strengths of our payload design is the originality of our autonomous real-time camera orientation system (ARTCOS). The other major strength can be found in the originality of our self-designed Printed Circuit Board layouts. This feature alone represents over 100 man hours of work. Along with space and power efficiencies, the PCB’s provide major enhancement of the signal integrity of the sensor data. For ease of reading, you will find documents such as itemized budgets, and launch procedures moved to the appendix along with Sensor and Material Safety Data sheets. We have enjoyed the challenges presented in the writing of this document and submit it for your review.
  2. 2. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review ii ii Table of Contents Table of Contents List of Figures..................................................................................................................vi List of Tables...................................................................................................................ix List of Acronyms..............................................................................................................xi I) Summary of PDR Report ........................................................................................... 13 Team Summary ......................................................................................................... 13 Launch Vehicle Summary .......................................................................................... 13 Payload Summary...................................................................................................... 13 Launch Vehicle Overview .......................................................................................... 14 Motor Selection....................................................................................................... 14 Recovery ................................................................................................................ 14 II) Changes Made Since Proposal................................................................................. 16 Vehicle Change log................................................................................................. 16 Payload Change Log .............................................................................................. 16 Project Plan Change Log........................................................................................ 17 III) Vehicle Criteria......................................................................................................... 18 Selection, Design, and Verification of Launch Vehicle ............................................... 18 Launch Vehicle Mission Statement......................................................................... 18 Mission Success Criteria ........................................................................................ 18 Propulsion Motor Selection .................................................................................... 26 Performance Characteristics and Verification Metrics................................................ 31 Recovery System.................................................................................................... 31 Structure System .................................................................................................... 32 Propulsion............................................................................................................... 32 Verification Plan...................................................................................................... 32 Risks and Plans for Reducing Risks .......................................................................... 39 Planning of Manufacturing ...................................................................................... 43 Confidence and Maturity of Design............................................................................ 45 Electrical Schematics of the Recovery System.......................................................... 47
  3. 3. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review iii iii Mass Statement......................................................................................................... 48 Recovery System....................................................................................................... 50 Recovery Component Itemization........................................................................... 61 Mission Performance Predic ...................................................................................... 66 Mission Performance Criteria.................................................................................. 66 Simulations ............................................................................................................. 66 Stability................................................................................................................... 71 Kinetic Energy......................................................................................................... 72 Interfaces and Integration .......................................................................................... 78 Internal Vehicle Interfaces ...................................................................................... 78 Vehicle to Ground Launch System Interfaces......................................................... 81 Launch Operation Procedures ................................................................................... 81 Safety and Environment (Vehicle).............................................................................. 81 The Safety Officer................................................................................................... 81 Failure Modes......................................................................................................... 82 Rocket Design Failure Modes................................................................................. 82 Payload Integration Failure Modes ......................................................................... 83 Launch Operations Failure Modes.......................................................................... 83 Hazard Analysis......................................................................................................... 85 Environment............................................................................................................... 88 Environmental effects of the project........................................................................ 88 Environmental effect on the project ........................................................................ 89 IV) Payload Criteria ....................................................................................................... 90 System Level Review ............................................................................................. 90 Required Subsystems................................................................................................ 94 Atmospheric Data Gathering .................................................................................. 94 Global Positioning System.................................................................................... 104 Wireless Transmitter............................................................................................. 105 Autonomous Camera Orientation System ............................................................ 106 Video Capture....................................................................................................... 111 Liquid Crystal Display ........................................................................................... 112
  4. 4. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review iv iv Official Scoring Altimeter ...................................................................................... 113 Power Supply........................................................................................................ 114 Performance Characteristics.................................................................................... 114 Verification Plan....................................................................................................... 115 Preliminary Integration Plan..................................................................................... 117 Precision of instrumentation, repeatability of measurement, and recovery system.. 118 Drawings and Electrical Schematics........................................................................ 120 Electrical Schematic................................................................................................. 121 Cross-Component Compatibility .............................................................................. 131 Payload Concept Features and Definition................................................................ 135 Uniqueness and Significance ............................................................................... 137 Suitable Level of Challenge .................................................................................. 137 Science Value.......................................................................................................... 138 Experimental Logic, Approach, and Method of Investigation ................................ 139 Relevance of Expected Data and Accuracy/Error Analysis .................................. 140 V) Project Plan ............................................................................................................ 140 Testing Budget...................................................................................................... 141 Outreach Budget................................................................................................... 141 Travel Budget ....................................................................................................... 142 Budget Summary.................................................................................................. 142 Funding Plan............................................................................................................ 143 Timeline ................................................................................................................... 143 Gantt Timeline ...................................................................................................... 143 Testing Gantt Timeline............................................................................................. 145 Outreach Gantt Timeline....................................................................................... 147 Outreach Plan ............................................................................................................. 148 Educational Outreach .............................................................................................. 148 Educator Outreach................................................................................................... 149 Community Outreach ............................................................................................... 149 Star Party.............................................................................................................. 149 Tarleton Regional Science Olympiad.................................................................... 149
  5. 5. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review v v Participation Goal..................................................................................................... 150 Accomplished Educational Outreach.................................................................... 150 Conclusion .................................................................................................................. 153 Appendix A - Itemized Subsystem Budget .................................................................. 155
  6. 6. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review vi vi List of Figures Figure 1 Vehicle Structure............................................................................................. 19 Figure 2 - Nose Cone.................................................................................................... 20 Figure 3 Ballast System ................................................................................................ 21 Figure 4 Upper Body Airframe....................................................................................... 22 Figure 5 Payload Housing............................................................................................. 24 Figure 6 - Fin Design Selection ..................................................................................... 25 Figure 7 - Fin Design Options ....................................................................................... 25 Figure 8 Selected Fin Design ........................................................................................ 26 Figure 9 – Cesaroni L1720 Thrust Curve ...................................................................... 28 Figure 10 - AeroTech L1390 Thrust Curve.................................................................... 29 Figure 11 - Cesaroni L1090 Thrust Curve..................................................................... 30 Figure 12 - Risk Plot...................................................................................................... 42 Figure 13 - Team Hierarchy .......................................................................................... 45 Figure 14 - Project Life Cycle ........................................................................................ 46 Figure 15 - Dimensional Drawings ................................................................................ 46 Figure 16 - Recovery Electrical Schematic.................................................................... 47 Figure 17 - Featherweight Raven 3 Wiring.................................................................... 48 Figure 18 - Flight Sequence.......................................................................................... 50 Figure 19 Daveyfire Electric Match................................................................................ 52 Figure 20 Altitude Simulation ........................................................................................ 67 Figure 21 Velocity Simulation........................................................................................ 68 Figure 22 Acceleration Before Burn Out Simulation...................................................... 69 Figure 23 Acceleration After Burn Out........................................................................... 70 Figure 24 L1720 Thrust Curve ...................................................................................... 71 Figure 25 Stability: Center of Pressure/Gravity ............................................................. 71 Figure 26 Weather Cocking........................................................................................... 76 Figure 27 Weather Cocking Simulation......................................................................... 77 Figure 28 Payload Pre-Integration ................................................................................ 78 Figure 29 Coupler Specifications .................................................................................. 79 Figure 30 Payload Design Configuration....................................................................... 91 Figure 31 Self-leveling Camera Configuration............................................................... 93 Figure 32 Arduino Mega 2560-R3 Microcontroller......................................................... 95 Figure 33 Adafruit Micro SD Adapter............................................................................. 97 Figure 34 BMP180 Breakout Board............................................................................... 98 Figure 35 MS5611-01BA03 Breakout Board................................................................. 98 Figure 36 HIH4030 Breakout Board ............................................................................ 100 Figure 37 HH10D Breakout Board .............................................................................. 100 Figure 38 SP-110 ........................................................................................................ 101 Figure 39 TAOS TSL2561 Breakout Board................................................................. 102
  7. 7. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review vii vii Figure 40 SU-100........................................................................................................ 103 Figure 41 TOCON_ABC3............................................................................................ 104 Figure 42 LS20031...................................................................................................... 105 Figure 43 XBee-PRO XSC S3B .................................................................................. 106 Figure 44 Arduino Pro Mini 328 Mircocontroller .......................................................... 107 Figure 45 VC0706 Photographic Camera ................................................................... 108 Figure 46 Servo Dimensional Layout .......................................................................... 109 Figure 47 ADXL345 Breakout Board........................................................................... 110 Figure 48 VCC-003-MUVI-BLK Video Camera ........................................................... 111 Figure 49 Sparkfun LCD-11062 Screen ...................................................................... 112 Figure 50 Adept A1E................................................................................................... 113 Figure 51 Atmospheric Data Gathering Conceptual Wiring......................................... 120 Figure 52 Camera Orientation Wiring.......................................................................... 121 Figure 53 PCB Schematic for Payload Sensors.......................................................... 122 Figure 54 I^2C Sensors............................................................................................... 123 Figure 55 Radio Communications ............................................................................... 124 Figure 56 MicroSD Card Reader................................................................................. 125 Figure 57 Analog Sensors........................................................................................... 126 Figure 58 USB Interface.............................................................................................. 127 Figure 59 GPS Module................................................................................................ 128 Figure 60 Voltage Regulators...................................................................................... 129 Figure 61 Preliminary PCB Layout .............................................................................. 130 Figure 62 Component Listing ...................................................................................... 131 Figure 63 SMT (left) vs. Through Hole Devices (right) ................................................ 131 Figure 64 Atmospheric Data Gathering Subsystem Data Flow ................................... 132 Figure 65 Autonomous Camera Orientation System Data Flow.................................. 133 Figure 66 Atmospheric Data Gathering Software Flow Chart...................................... 134 Figure 67 Autonomous Camera Orientation Software Flow Chart .............................. 135 Figure 68 - PCB Board................................................................................................ 136 Figure 69 - Self-Leveling Camera System................................................................... 137 Figure 70 - Initial Funding............................................................................................ 143 Figure 71 Project Timeline .......................................................................................... 144 Figure 72 Testing Timeline.......................................................................................... 146 Figure 73 Outreach Timeline....................................................................................... 147 Figure 74 - Acton Middle School ................................................................................. 148 Figure 75 Subject Interest ........................................................................................... 151 Figure 76 Presentation Learning Outcomes................................................................ 152 Figure 77 Favorite Part................................................................................................ 153
  8. 8. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review viii viii
  9. 9. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review ix ix List of Tables Table 1 Launch Vehicle Summary ................................................................................ 13 Table 2 Payload Summary............................................................................................ 13 Table 3 Vehicle and Recovery Changes ....................................................................... 16 Table 4 Payload Change............................................................................................... 17 Table 5 Project Plan Changes....................................................................................... 17 Table 6 Upper Body Airframe Trade and Selection....................................................... 21 Table 7 Pros and Cons of Payload Options .................................................................. 23 Table 8 Motor Trade and Selection ............................................................................... 27 Table 9 Vehicle Verification Table................................................................................. 37 Table 10 Recovery System Verification Table............................................................... 39 Table 11 System Risks.................................................................................................. 41 Table 12 Project Risks .................................................................................................. 42 Table 13 Testing Summary ........................................................................................... 43 Table 14 Total Mass Summary ..................................................................................... 48 Table 15 Payload Mass Summary ................................................................................ 49 Table 16 Recovery Mass Summary .............................................................................. 49 Table 17 Structure Mass Summary............................................................................... 50 Table 18 Deployment Altimeter Trade and Selection .................................................... 51 Table 19 Electric Match Trade & Selection ................................................................... 53 Table 20 Recovery Testing Dates ................................................................................. 56 Table 21 Anemometer Trade & Selection ..................................................................... 56 Table 22 Parachute Diameters...................................................................................... 61 Table 23 Main Parachute Trade and Selection ............................................................. 62 Table 24 Drogue Parachute Trade and Selection ......................................................... 62 Table 25 Parachute Protection Materials ...................................................................... 63 Table 26 Shock Chord Trade and Selection.................................................................. 64 Table 27 Launch Day Recovery System Budget........................................................... 66 Table 28 Kinetic Energy Summarization ....................................................................... 74 Table 29 Landing Radius .............................................................................................. 77 Table 30 Ground- Vehicle Interface .............................................................................. 81 Table 31 Potential Failure Modes for the Design of the Vehicle.................................... 83 Table 32 Potential Failure Modes during Payload Integration....................................... 83 Table 33 Potential Failure Modes during Launch .......................................................... 85 Table 34 Potential Hazards to Personnel...................................................................... 87 Table 35 Summary of Legal Risks ................................................................................ 88 Table 36 Effects of Materials used in Construction and Launch.................................... 89 Table 37 Environmental Factors ................................................................................... 90 Table 38 Top Four Candidates for Microcontroller Selection ........................................ 96 Table 39 Top Four Choices for Data Storage Medium.................................................. 96
  10. 10. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review x x Table 40 Micro SD Adapter Selection ........................................................................... 97 Table 41 Pressure Sensor Trade and Selection............................................................ 99 Table 42 Top Considerations for Temperature Sensor Selection.................................. 99 Table 43 Humidity Sensors Trade and Selection ........................................................ 101 Table 44 Pyranometers Trade and Selection .............................................................. 102 Table 45 UV Sensor Trade and Selection................................................................... 104 Table 46 GPS Modules Trade and Selection .............................................................. 105 Table 47 XBee Module Trade and Selection............................................................... 106 Table 48 Microcontrollers Trade and Selection ........................................................... 108 Table 49 Camera Modules Trade and Selection ......................................................... 109 Table 50 Servo Motors Trade and Selection............................................................... 110 Table 51 Accelerometers Trade and Selection ........................................................... 111 Table 52 Top Two Video Cameras Trade and Selection............................................. 112 Table 53 LCD Screen Trade and Selection................................................................. 113 Table 54 Official Scoring Altimeters Trade and Selection ........................................... 114 Table 55 Batteries Trade and Selection ...................................................................... 114 Table 56 Payload Subsystems Evaluation and Verification Metrics ............................ 115 Table 57 SOW Verification.......................................................................................... 117 Table 58 Payload Sensors Precision........................................................................... 119 Table 59 Payload Objectives Summary ...................................................................... 139 Table 60 Proposed Rocket Vehicle Budget Summary................................................. 141 Table 61 Preliminary Testing Budget Summary .......................................................... 141 Table 62 Outreach Budget .......................................................................................... 142 Table 63 Estimated Travel Budget .............................................................................. 142 Table 64 Preliminary Budget Summary....................................................................... 143 Table 65 Accomplished Educational Outreach............................................................ 150 Table 66 Presentation Learning Outcomes................................................................. 151 Table 67 Favorite Part................................................................................................. 152 Table 68 Preliminary Structure Budget............................ Error! Bookmark not defined. Table 69 Recovery System Budget................................. Error! Bookmark not defined. Table 70 Payload Budget................................................ Error! Bookmark not defined.
  11. 11. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review xi xi List of Acronyms ADC-Analog to Digital Converter AGL-Above Ground Level APCP-Ammonium Perchlorate Composite Propellant APR-Automatic Packet Reporting System BMP-Barometric Pressure BOD-Board of Directors CAD-Computer Aided Drafting CFR-Critical Design Review CNC-Computer Numerical Control DC-Direct Current EEPROM-Electrically Erasable Programmable Read-Only Memory EMF-Electric and Magnetic Fields FAA-Federal Aviation Administration FRR-Flight Readiness Review GGA-Global Positioning System Fix Data GLL-Graphic Position-Latitude/Longitude GNSS-Global Navigation Satellite System GPS-Global Positioning System GSA-GNSS and Active Satellites GSV-GNSS Satellites in View HCI-Harris Composites Incorporated LCD-Liquid Crystal Display LCS-The Launch Control System LED-Light Emitting Diode LSO-Launch Safety Officer MHz-Mega Hertz MSDS-Material Safety Data Sheet N/A-Not Available NAR-National Association of Rocketry NASA-National Aeronautics and Space Administration NMEA-National Marine Electronics Association NMEA-National Marine Electronics Association NOAA-National Oceanic and Atmospheric Administration OSHA-Occupation Safety and Health Administration PCB-Perforated circuit board RF-Radio Frequency RMC-Recommended Minimum Specific GNSS Data RSO-Range Safety Officer SD-Secure Digital SLP-Student Launch Projects SMD-Science Mission Directorate SOW-Statement of Work SPI-Serial Peripheral Interface STEM-Science Mathematics Engineering Technology
  12. 12. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review xii xii TAP-Technical Advisor Panel TRA-Tripoli Rocketry Association TTL Transistor-transistor logic USLI-University Student Launch Initiative UV-Ultraviolet
  13. 13. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 13 I) Summary of PDR Report Team Summary Tarleton Aeronautical Team Tarleton State University Box T-0470 Stephenville, Texas 76402 Team Mentor Our team mentor is Pat Gordzelik. TRA 5746, L3. TAP NAR 70807 L3CC Committee Chair. Founder/past Prefect/President Potrocs (Panhandle of Texas Rocketry Society) Inc., Tripoli Amarillo # 92. BOD member, Tripoli Rocketry Association Inc. Married to Lauretta Gordzelik, TRA 7217, L2. Launch Vehicle Summary Size and Mass Length 108 inches Outer Diameter 5.525 inches Mass 33.5 pounds Motor Selection Cesaroni L1720-WT-P Recovery Drogue 24” Nylon Parachute, Apogee Deployment Main 120” Nylon Parachute, 500 foot AGL Deployment Avionics Primary Featherweight Raven 3 Altimeter, Backup PerfectFlite Stratologger Altimeter, and Beeline GPS Tracking Table 1 Launch Vehicle Summary The Milestone Review Flysheet can be found in Appendix L. Payload Summary Title Experiment Science Mission Directorate Payload Gather Atmospheric and GPS Data, Autonomously Orientate Photographic Camera, Capture Video for Public Outreach, and a Clear Acrylic Housing Table 2 Payload Summary
  14. 14. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 14 Launch Vehicle Overview Size and Mass The total rocket body is 108 inches in length and has a diameter of 5.5 inches. There are four body sections: a nose cone, upper body airframe, payload, and a booster section. The nose cone is fiberglass, consists of an elliptical shape, and has a 7.5 inch length with a 5.5 inch shoulder. The upper body airframe is 28 inches in length and consists of fiberglass. The payload is 36 inches in length and is clear acrylic. The booster section is 36 inches in length and consists of fiberglass. The rocket has a four fin configuration; each fin has a height of five inches. The simulated unballasted mass of the entire rocket is approximately 33.5 pounds. Motor Selection The chosen motor is a Cesaroni L1720-WT-P. The motor has an average thrust of 1754 Newtons and a maximum thrust of 1947 Newtons. The total impulse is 3,696 Newton seconds. It has an initial launch weight of 7.37 pounds and a post burn weight of 3.5 pounds. This motor is chosen based on simulations taking into account average launch conditions of the launch site and date. The high initial thrust is also a factor. The high thrust provides a launch rail exit velocity of 75 feet per second. With average conditions simulated and no ballast, the predicted vehicle apogee is 5,342 feet. This apogee, slightly above one mile, allows for the total mass to increase. Recovery The upper body airframe comprises of the main parachute compartment and the main altimeter compartment. The booster sections include the drogue parachute compartment and the drogue altimeter compartment. The drogue parachute deploys at apogee and the main parachute deploys at 500 feet above ground level. Each deployment altitude is monitored by a redundant system of altimeters, each is capable of being magnetically locked in the “on” position for the duration of the flight. Once apogee is detected, an electronic match charges to ignite a black powder well which will cause ejection of the drogue parachute. The same mechanism ejects the main parachute. Both parachutes shield from the heat of the ejection charges by Nomex parachute bags. Shock harnesses keep the four body sections tethered after both parachute ejections. Parachute sizes and deployment heights calculate to ensure that launch vehicle recovery occurs within a 2500 foot radius of
  15. 15. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 15 the launch site. These calculations also ensure that the sections have a maximum kinetic energy of less than 75 foot pounds force. Science Mission Directorate Payload The payload, at a minimum, fulfills the requirements of the SMD payload. This includes storing and transmitting atmospheric sensor measurements, GPS data, and taking pictures of the proper orientation. The payload contains the necessary atmospheric sensors and an autonomous camera orientation system. An onboard LCD screen assists in pre-launch operations. The payload houses a video camera to store footage of the flight for educational engagement and public relations purposes. A clear, acrylic, tube houses the SMD payload. Durable aluminum mounting rails secure components. The clear acrylic housing allows for internal solar data gathering, pictures to be taken from within, and visual inspection of payload components.
  16. 16. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 16 II) Changes Made Since Proposal In the following sections, tables outline the changes since the proposal. These reflect outstanding design conflicts that require change in order to meet the criterion of the project. Vehicle Change log As the vehicle design evolves, changes take place and documentation is performed in order to accommodate healthy communication on the new designs. The following chart describes the changes thus far in the vehicle design. Vehicle and Recovery Changes Rationale Maximum outer diameter from 4 inches to 5.525 inches Payload section integration and design is better suited for the wider diameter Total length from 103 in to 108 in To allow some additional room for recovery system implementation Fin size and design To maintain stability for a larger vehicle Incorporation of a ballast system in the nose cone To allow a greater level of ballast weight control on launch day Nose cone length from 7 inches to 7.5 inches Commercial availability after diameter change Black powder charges more powerful Volume increase within recovery system compartments Wire mesh added to recovery altimeter compartments To provide shielding from wireless transmitting devices Table 3 Vehicle and Recovery Changes Payload Change Log As the current payload design progresses for engineering efficiency, a change log is kept in order to inform members of changes and facilitate communication. The following chart recognizes changes pertaining to the payload design. Payload Changes Reason Payload width from 3.5” to 5" To provide more room for autonomous camera orientation system and inclusion of a video camera Payload housing will be mounted by screws through the acrylic to a fiberglass coupler Avoids manufacturing a threaded acrylic cap
  17. 17. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 17 Inclusion of servos and an accelerometer for the autonomous camera orientation system Critical payload objective PCB planning and design To reduce overall space required as compared to perforated board mounting The official scoring altimeter moved from recovery bays to payload housing To avert pressure fluctuations from ejection charges, to be seen through the clear acrylic Inclusion of video camera For public relations and educational outreach Number of humidity sensors from 1 to 2 For redundancy in humidity measurements Number of pressure sensors from 4 to 2 For efficiency of design Number of pyranometers from 2 to 4 180 Field of view Number of UV sensors from 2 to 4 180 Field of view The top of the payload rail system will rest in a milled grove of the upper payload mounting cap For structural support Two, 90 degree opposing boards for solar irradiance and UV sensors Solar irradiance and UV sensors can gather data at 90 degree intervals around the rocket vehicle Handheld Yagi directional antenna will be used to receive all transmitted data from the payload Development of an automated tracking system for the ground station antenna was deemed unnecessary Table 4 Payload Change Project Plan Change Log As the project develops a change log shall be kept in order to effectively communicate to the team as a whole the modifications to the project plan. The chart that follows includes the list of known changes thus far to the project plan. Project Plan Changes Rationale Testing Plan and Timeline has been further developed Time Constraints Educational Outreach Plan has been further developed To increase quality of Educational outreach aspect to the entire project Community Outreach events added To increase quality of public outreach and project sustainability Table 5 Project Plan Changes
  18. 18. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 18 III) Vehicle Criteria Selection, Design, and Verification of Launch Vehicle Launch Vehicle Mission Statement The mission is to design, build, and launch a reusable vehicle capable of delivering a payload to 5,280 feet above ground level (AGL). The vehicle will carry a barometric altimeter for official scoring as well as the Science Mission Directorate (SMD) payload. The design of the vehicle ensures a subsonic flight and must be recoverable and reusable on the day of the official launch. The launch vehicle meets the customer prescribed requirements set forth in the Statement of Work (SOW) of the NASA 2012-2013 Student Launch Projects (SLP) handbook. Launch Vehicle Requirements The vehicle adheres to the following primary requirements. The complete list of requirements is in the Vehicle Verification Table Table 9.  Vehicle shall carry a scientific or engineering payload. (Requirement 1.1)  Vehicle shall reach an apogee altitude of 1 mile AGL. (Requirement 1.1)  Vehicle shall carry one official scoring altimeter. (Requirement 1.2)  Vehicle must remain subsonic from launch until landing. (Requirement 1.3)  Vehicle must be recoverable from a 2500 foot radius away from the launch pad and reusable on the day of the official launch. (Requirement 2.3)  Vehicle must use a commercially available APCP motor with no more than 5,120 Newton-seconds of impulse. (Requirement 1.11, 1.12) Mission Success Criteria The project defines a successful mission as a flight with payload, where the vehicle and SMD payload are recovered and able to be reused on the day of the official launch. Moreover, the vehicle will not exceed 5,600 feet and the official scoring altimeter will be intact and report the official altitude. The recovery system stages a deployment of the drogue parachute at apogee and follows deployment of the main parachute at 500 feet. After apogee and descent, the entire vehicle lands within 2,500 feet of the launch pad. System Level Review This section reviews the design of the vehicle which includes structure, propulsion, and recovery. The project requires the team to consider, research, and analyze
  19. 19. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 19 various concepts. Each system contributes to the overall success of the mission and helps determine how useful each system is. Trade and selection is performed for the structure and propulsion systems. The Recovery Subsystem contains recovery, trade, and selection. Calculations and measurements for each system are a part of the presentation. Structure Selection of vehicle components contains variables of durability, aerodynamics, cost, availability, and functionality. The structure of the vehicle composes of the nose cone, upper body/lower body airframe, payload section, tube coupling, and fin structure. The structure is able to withstand the substantial forces throughout the flight. It must also remain subsonic and be reusable. Material, size, and design must be adequately chosen to fulfill many requirements. The material composition is chosen primarily to remain reusable. The elliptical nose cone, four fin design, and material composition of the rocket all contribute to remaining subsonic throughout the entire flight. The material composition of the payload section is chosen to be transparent. The fins are also chosen for their aerodynamic properties. The diameter must be able to house the SMD payload and parachute deployment systems. The coupling system was designed to couple two different sizes of inside diameters and to ensure separation or non-separation. Figure 1 displays the structure of the vehicle. Figure 1 Vehicle Structure Nose Cone The design is an elliptical nose cone. The vehicle is aerodynamic and remains
  20. 20. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 20 subsonic as per the requirements for the flight. The four types of nose cones for considerations are elliptical, parabolic, conical, and Von Karman. At subsonic speeds, elliptical and parabolic nose cones generate the least amount of drag. The conical and Von Karman style nose cones are longer and add excess weight and more drag at subsonic speeds, so they are not well fit for this project. The parabolic nose cone experiences slightly less drag but has slightly more weight in comparison to the elliptical nose cone. An elliptical nose cone is ideal due to the commercial availability and manufacturing cost of a parabolic nose cone. Figure 2 - Nose Cone The nose cone in Figure 2 above is 13 inches long including the 5.5 inch shoulder, and is 0.075 inch thick. These dimensions accommodate the ballast system, seen in Figure 3. The ballast system consists of two bulkheads, one higher in the nose cone that is not removable and one at the exit of the shoulder that is removable. A five inch bolt stretches between the bulkheads and washers are added to adjust the ballast mass.
  21. 21. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 21 Figure 3 Ballast System Upper Body Airframe/Booster Section Fiberglass is the material of choice for the upper body airframe and booster section. The durability of fiberglass improves the chances of the rocket being reusable (Requirement 1.4). Fiberglass is also readily available. The upper body airframe houses the main parachute deployment system. The lower body airframe houses the drogue parachute deployment system, the motor mount, and secures the fins. These components need to withstand all stresses present during flight. Three materials are considerations for the upper and lower body airframes: blue tube 2.0, fiberglass, and glass phenolic tubing. Table 6 represents a trade and selection of the component materials in the upper body airframe. Material Peak Load (lbf) Peak Stress (psi) Modulus Avg. Cost per Foot Availability (1 highest) Fiberglass 19256.1 37806.2 2980.8 $38.84 1 Blue Tube 3211.1 5293.4 607.1 $13.74 3 Glassed Phenolic 7758.9 8983.4 1228.9 $50.27 2 Table 6 Upper Body Airframe Trade and Selection
  22. 22. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 22 The upper body airframe is 28 inches long and the lower body airframe is 36 inches long. Both are 0.075 inches thick. These lengths provide adequate space for the deployment systems. Figure 4 depicts the upper body airframe with nose. The booster Figure 4 Upper Body Airframe Payload Structure The payload housing structure consists of clear acrylic. This section houses the SMD payload. Considerations exist for three options for the payload housing structure. Option one was an independent fiberglass section. Option two consists of an extension of the upper body airframe. Option three is an independent acrylic section. Payload Airframe Options Separate Fiberglass Airframe Extended Upper Airframe Acrylic Airframe
  23. 23. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 23 The first option is to couple the payload housing structure to the upper airframe. The payload is removable in order to inspect electronics. The sensors for detecting ultraviolet radiation and solar irradiance have to be mounted on the exterior of the vehicle. The second option is to have one upper body section that contains the main parachute deployment system in the first half and the SMD payload in the second half. This reduces the amount of weight because of the lack of a second coupler. Sensors would have to mount on the exterior and electronics checked through dismantling. The third option is an acrylic payload section. This allows all sensors and the camera to mount internally. Also, this allows us to visually check that all electronics are functional. Acrylic is also chosen because of its similar properties to fiberglass. This option does not require the payload to be deployable from the vehicle to gather data. This is safer and eliminates the chance of failure upon ejection of the payload. The material selected for the payload housing structure is cast acrylic. This is the best option for the electronics inspection, camera and sensor functionality and safety. The payload section, seen in Figure 5, is 36 inches long, 5.5 inches wide, and 0.125 inches thick. The increase in thickness of the acrylic is chosen to ensure adequate strength for the payload section. The compressive properties of acrylic are undergoing testing with standard ASTM methods to verify the integrity of this selection. The acrylic tube manufacturer provides a data sheet on this selection (refer to Appendix K). The compressive yield strength of acrylic, according to MatWeb (Material Property Data website), is 18,000 psi. The force the rocket Pros Durability, cost, availability, familiarity Durability, cost, availability, familiarity, lack of coupling system (lighter) Visibility of electronics and payload components from exterior, internally mounted sensors Cons Inability of external inspection, externally mounted sensors, coupling system Inability of external inspection, externally mounted sensors, inaccessibility to internal systems Durability, weight Table 7 Pros and Cons of Payload Options
  24. 24. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 24 motor exerts is 386 pounds and the weight of the upper body airframe and nose cone is 7.42 pounds. Therefore, the vehicle is able to withstand loading during the flight. Additional testing at a materials testing facility, as well as full scale pending launch testing provide more accurate data on the acrylic payload. Figure 5 Payload Housing Airframe Coupling The coupling system must keep vehicle sections together during flight and separate as needed. The vehicle design requires coupling the upper body airframe to the acrylic payload structure. These have different inside diameters. The design provides for a single coupler to fit both. It is 11 inches long, allowing 5.5 inches of insertion into each section. The outside diameter of the coupler is 5.25 inches to meet the acrylic inside diameter. The inside diameter of the fiberglass upper airframe is 5.375 inches. Thus, 5.5 inches of the coupler wrap in fiberglass and resin to increase the outside diameter to 5.375 inches. Since the upper body section does not need to separate from the payload section, these sections are rivet to the coupler. Rivets are superior to epoxy in fastening and allow the area inside the coupler to be accessible by a user by removing the rivets. The lower body sections are friction fitted to allow separation with the black powder charges. This coupler is 11 inches in length, allowing 5.5 inches of insertion into each section. Fin Structure Figure 6 is an exact dimensional drawing of the fin design selection.
  25. 25. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 25 Figure 6 - Fin Design Selection The fins are the most important component to flight stability. They determine the center of pressure, create drag, and create the corrective moment force for stability. The primary option for fins is three versus four fins. Three fins provide less drag, less weight, and less corrective moment force. The change in stability margin between three fins and four fins is insignificant because the weight added moves back the center of pressure and center of gravity nearly equal amounts. Although the four fins weigh more, they create more corrective moment force for stability. The four fins also create more drag on the vehicle, which is desired for the motor choice. Using Open Rocket to simulate the vehicle design, the fin design simulations rend these the findings in Figure 7: Alternative Option 2Alternative Option 1Selected Option Figure 7 - Fin Design Options
  26. 26. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 26 The selected option achieves proper altitude and provides a good stability margin. The alternative option one performs well, but the reusability of the vehicle could be compromised. This option could result in structural damage upon landing, as the fins extend beyond the base of the vehicle. Alternative option two increases drag, resulting in the vehicle failing to reach altitude with the selected motor. According to data from simulations, the selected option is chosen in a four fin design as seen in Figure 8. Root Chord: 12 inches Tip Chord: 0 inches Height: 5 inches Sweep Length: 9.8 in Sweep Angle: 63 degrees Propulsion Motor Selection The selected motor is a Cesaroni L1720-WT-P. The high initial thrust helps to stabilize the rocket as it departs from the launch rail. Its high thrust to weight ratio is also beneficial to stability. Through simulations that take into consideration the average conditions for the launch site and date, the Cesaroni L1720-WT-P causes the vehicle to achieve an apogee of one mile AGL. A ballast system alters the rocket’s weight to account for alterations between conditions on launch date and simulated conditions to ensure apogee height of one mile AGL. The three viable options for motors are Aerotech L1390G, Cesaroni L1090SS-P, and Cesaroni L1720-WT-P. Each of the motors has a diameter of 2.95 inches. OpenRocket simulates all motor options. Motors options must achieve apogee above one mile allowing for the incorporation of ballast weight as well as the increase to vehicle mass as the design evolves. The Aerotech L1390G motor results in an apogee height of 5618 feet and off the rail velocity of 65.6 feet per second, which is acceptable but not optimal. The Cesaroni L1720-WT-P achieved just over one mile, which allows for weight increase and ballast if needed. Its Figure 8 Selected Fin Design
  27. 27. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 27 velocity off the rail is 75.9 feet per second, providing adequate stability. The Cesaroni L1090SS-P achieves an apogee of 1200 feet above one mile, and a fully ballasted configuration will not achieve the desired altitude. Table 8 is a trade and selection table of the motor options. Motor Apoge e (ft.) Velocity Off Rail (ft./s) Total Impulse Max. Velocity (ft./s) Average Thrust Burn Time (s) Thrust to Weight Ratio Availability / Cost Cesaroni L1720- WT-P 5345 75.9 830.9lbf s (3696Ns ) 738 394.3lbf (1754N) 2.15 11.8 High/ $170.96 Aerotech L1390G 5618 65.6 887.8lbf s (3949Ns ) 723 308.9lbf (1374N) 2.65 9.2 Medium/ $209.99 Cesaroni L1090SS- P 6479 76.8 1082lbfs (4815Ns ) 733 246.6lbf (1097N) 4.4 7.4 Medium/ $346.95 Table 8 Motor Trade and Selection Cesaroni L1720 The Cesaroni L1720 has a total impulse of 3696 Newton-seconds, which does not exceed the total impulse maximum of 5120 Newton-seconds. The motor’s corresponding thrust curve as calculated by Rocksim software is represented in Figure 9. As shown in the thrust curve, the motor has a fairly neutral motor burn. As shown in Table 8 and marked in Figure 9, average thrust for this motor is 394.3lbf = 1754N. With this motor, the launch mass of the rocket is 536oz = 15.2kg. Noting that the acceleration of gravity is approximately 9.8m/s², for this motor the thrust to weight ratio is achievable by = 11.8 : 1, which exceeds the suggested ratio of 5 : 1.
  28. 28. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 28 Figure 9 – Cesaroni L1720 Thrust Curve Aerotech L1390G The Aerotech L1390G has a total impulse of 3949 Newton-seconds, which does not exceed the total impulse maximum of 5120 Newton-seconds. The motor’s corresponding thrust curve as calculated by the Rocksim software is represented in Figure 10. As shown in the thrust curve, the motor begins with a progressive motor burn reaches maximum thrust and then begins a regressive motor burn. As shown in the above table and marked in thrust curve, average thrust for this motor is 308.9lbf = 1374N. In the following calculation, the mass of the rocket at launch is used because it represents the maximum mass that the motor would have to be in order to lift during the flight. In order that the motor be able to lift the rocket, it must produce enough thrust to overcome the force of gravity, or enough mechanical energy to achieve a thrust to weight ratio of at least 1.0. In general for a high- powered rocket, the thrust to weight ratio is given by ( ) ( ) .
  29. 29. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 29 With this motor, the launch mass of the rocket is 557.3oz = 15.8kg. Noting that the acceleration of gravity is approximately 9.8m/s², for this motor the thrust to weight ratio is given by = 9.2 : 1, which exceeds the suggested ratio of 5 : 1. Figure 10 - AeroTech L1390 Thrust Curve Cesaroni L1090 The Cesaroni L1090 has a total impulse is 4815 Newton-seconds, which does not exceed the total impulse maximum of 5120 Newton-seconds. The motor’s corresponding thrust curve as calculated by the open rocket software is represented in Figure 11. As shown in the thrust curve, the motor quickly reaches the maximum thrust then starts a regressive motor burn. As shown in Table 8 on page 27 and marked in Figure 11, average thrust for this motor is 246.6lbf = 1097N. With this motor, the launch mass of the rocket is 610.2oz = 17.3kg. Noting that the acceleration of gravity is approximately 9.8m/s², for this motor the thrust to weight ratio is given by = 6.47 : 1, which exceeds the suggested ratio of 5 : 1.
  30. 30. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 30 Figure 11 - Cesaroni L1090 Thrust Curve Motor Retainer The motor retainer needs to withstand the massive force of the motor on the rocket. Three commercially available motor retainers are considerable options. The first is a quick release system using a cap that snaps onto the retainer body. The second option is the same retainer body design, only with the implementation of a threaded cap. Both of these options are simply glued to motor tube. The third option includes a flange around the retainer body that allows for 12 screws to mount the retainer to the lower centering ring of the motor tube. It is also glued to the motor tube and uses a threaded cap for securing the motor in place. This appears to be the optimal option because of the added protection of mounting to the centering ring. Recovery System The recovery subsystem contributes to the overall mission by ensuring that the vehicle lands in a reusable condition within a 2500 foot radius from the launch site. Landing in a completely reusable condition entails that the payload and all other electronic and mechanical components remain in sound condition throughout the flight, including impact. To ensure that no component sustains irreparable damage
  31. 31. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 31 during ascent, the recovery electronics are mechanically armed and programmed to eject the drogue and main parachutes at safe altitudes. The ejection charges are sealed off and triggered electronically by redundant altimeter systems. All bulkheads and materials acted on by the charge firing are shown through testing to be sufficiently durable to withstand both the energetic impact and the heat of separation and ejection. Ensuring that the payload remain functional as a result of impact allows for the continual transmission of data upon landing, some of which will aid in physically locating the vehicle for a full recovery. Performance Characteristics and Verification Metrics Recovery System The performance of the recovery system relates to the vehicle’s ability to safely return to the ground. The recovery system must manage the speed of the vehicle in order to keep the kinetic energy of the each section below 75 ft-lbf. A kinetic energy greater than 75 ft-lbf could result in the structural failure and cause the vehicle to be non-reusable. To perform evaluation of parachute size, these equations are necessary: ∑ These equations are useful during multiple full scale test launches. Performance of the recovery system also depends on the correct operation of the deployment altimeters. Necessary evaluation of the recovery altimeters comes through research and design of electrical wiring diagrams. Verification of performance comes through ground testing and test launches.
  32. 32. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 32 Structure System The performance of the structure system is characterized by the ability to effectively allow the vehicle to be reusable and to efficiently integrate each subsystem. In the structure design are the materials in use and fin design, which contribute to the flight stability of the vehicle. The performance of the fins affects the entire rocket in that if they fail, the flight becomes unstable and possibly unsuccessful. Unsuccessful material choice or integration leads to failure if the rocket cannot withstand the forces of the motor upon launch and the force of impact with the ground upon landing. Materials’ evaluation is achievable through material properties databases and the undertaking of tests to withstand the force of the motor. It withstands testing from the Harris Composite Inc. material testing facility. The selected materials are then fully verifiable through the full scale test launches. Propulsion The performance characteristic of the propulsion subsystem lies in how consistently the motor performs so that flight predictions calculate accurately. Research and simulation achieve the evaluation of the propulsion system. The motor retainers must withstand the force of the motor, and failure results in entire mission failure. Both the motor and the retainer undergo static testing to ensure accuracy and strength. Full scale test launches provide data for verification. Trade and selection is viewable in Table 8. Verification Plan The verification plan in effect reflects how each requirement to the vehicle and recovery system satisfies its function. Requirements from the SOW are listed and paraphrased, followed by the satisfying feature of the design to that requirement. Ultimately, each design feature undergoes verification to ensure that it actually meets its requirements. Testing, analysis, and inspection serves as the mode of verification for each feature. A detailed Gantt chart containing test dates is in Figure 62.
  33. 33. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 33 Requirement (SOW) Vehicle Requirement Satisfying Design Feature Verification Method 1.1 Vehicle shall deliver payload to 5,280 feet AGL Motor selection Testing, Analysis 1.2 Vehicle shall carry one official scoring barometric altimeter Altimeter Model X is used and included in the SMD payload section Inspection 1.2.1 Official scoring altimeter shall report the official competition altitude via a series of beeps Altimeter Model X has this functionality Testing, Inspection 1.2.2 Teams may have additional altimeters Four additional altimeters, outside of the payload, will be used to detect apogee and ignite ejection charges Inspection 1.2.2.1 At Launch Readiness Review, a NASA official will mark the altimeter to be used for scoring Official altimeter placement allows ease of locating and marking Inspection 1.2.2.2 At launch field, a NASA official will obtain altitude by listening to beeps reported by altimeter The official altimeter has this functionality Testing, Inspection 1.2.2.3 At launch field, all audible electronics except for scoring altimeter shall be capable to turn off No other electronics in the design have audible indicators Inspection
  34. 34. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 34 1.2.3.1 Official, marked altimeter is damaged and/or does not report an altitude with a series of beeps Functional Recovery System Testing, Inspection 1.2.3.2 Team does not report to NASA official designated to record altitude with official marked altimeter on launch day This task will be assigned to an appropriate team member Analysis, Inspection 1.2.3.3 Altimeter reports apogee altitude of over 5,600 feet Motor selection Testing, Analysis 1.3 Launch vehicle remains subsonic from launch until landing Motor selection Testing, Analysis 1.4 Vehicle must be recoverable and reusable Recovery system allows a safe landing of vehicle Testing, Inspection, Analysis 1.5 Launch vehicle shall have a maximum of four independent sections Vehicle is composed of 3 tethered sections Inspection 1.6 Launch vehicle shall be prepared for flight at launch site within 2 hours Launch operations and assembly procedure Testing, Inspection
  35. 35. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 35 1.7 Launch vehicle will remain launch-ready for a minimum of one hour with critical functionality All critical on-board components will have sufficient capacity to meet entire system runtime (1.5 hours) Testing, Inspection, Analysis 1.8 Vehicle shall be compatible with either 8 feet long 1 inch rail (1010) 1010 rail buttons attached to vehicle body Inspection 1.9 Launch vehicle will be launched by a standard 12 volt DC firing system Motor selection is compatible with this firing system Inspection 1.10 Launch vehicle shall require no external circuitry or special equipment to initiate launch Motor ignition only requires the 12V DC firing system Inspection 1.11 Launch vehicle shall use a commercially available, certified APCP motor Cesaroni L1720 Inspection 1.12 Total impulse provided by launch vehicle will not exceed 5,120 Newton-seconds 3695.6 Ns Inspection 1.15 The full scale rocket, in final flight configuration, must be successfully launched and recovered prior to FRR Testing Schedule Testing 1.15.1 Vehicle and recovery system function as intended Featherweight and Stratologger Altimeters, Parachute Calculations Inspection 1.15.2 Payload does not have to be flown during full-scale test flight. The schedule allows for the payload to be flown in the full scale launch. Schedule
  36. 36. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 36 1.15.2.1 If payload is not flown, mass simulators shall be used to simulate payload mass If this occurs mass will be added proper sections. Testing 1.15.2.1.1 Mass simulators shall be located in same location on rocket as the missing payload mass Mass of each section is calculated, and mass will be added in the appropriate sections. Testing 1.15.2.2 Any energy management system or external changes to the surface of the rocket shall be active in full scale flight There will be no changes to the external surface of the rocket. Design 1.15.2.3 Unmanned aerial vehicles, and/or recovery systems that control flight path of vehicle, will fly as designed during full scale demonstration flight N/A N/A 1.15.3 Full scale motor does not have to be flown during full scale test flight The schedule and budget plan for the full scale motor to be flown. Schedule 1.15.4 Vehicle shall be flown in fully ballasted configuration during full scale test flight The schedule plan for the fully ballasted system. Schedule 1.15.5 Success of full scale demonstration flight shall be documented on flight certification form, by a Level 2 or Level 3 NAR/TRA observer, and documented in FRR package Pat Gordezlick will be present at the full scale launch. Schedule 1.15.6 After successfully completing full- scale demonstration flight, launch vehicle or any components shall not be modified without concurrence of the NASA Range Safety Officer (RSO) The schedule plans for the design to be complete and changes to be complete. Schedule 1.16 Maximum amount teams may spend on rocket and payload is $5000 Budget indicates that the total spent on the rocket and payload is less than $5000. Inspection, Analysis
  37. 37. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 37 Table 9 Vehicle Verification Table Table 10 is the verification table for the recovery system. Requirement (SOW) Recovery System Requirement Satisfying Design Feature Verification Method 2.1 Recovery devices shall be staged such that a drogue parachute is deployed at apogee; main parachute is at 500ft. Altimeters will stage ejection charges for respective parachutes at the prescribed altitudes Testing 2.2 Each independent section of the launch vehicle will have a maximum KE of 75 ft-lbf. Main parachute selection Testing, Analysis 2.3 Each independent section of the vehicle shall land with 2500 ft. of the launch pad Drogue and main parachute selection Testing, Analysis 2.4 Recovery electrical circuits shall be independent of payload electronics Recovery circuits are independent of payload with dedicated power supplies Inspection 2.5 Recovery system must include redundant altimeters Each deployment event is controlled by a main and backup altimeter Inspection 2.6 Each altimeter shall be armed in launch configuration with external arming switches Port holes in vehicle airframe to altimeter bays Inspection
  38. 38. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 38 2.7 Each altimeter shall have a dedicated power supply Each altimeter uses a separate 9 Volt battery Inspection 2.8 Each arming switch can be locked in the ON position Port holes in vehicle airframe to altimeter bays Testing, Inspection 2.9 Each arming switch will be a max of 6ft above the vehicle base Main altimeter bay is located at 6 feet above the base; drogue altimeter bay is located at 2 feet 4.5 inches from the base Inspection 2. 10 Removable shear pins shall be used for the drogue and main parachutes compartments Nylon shear pins will be used to couple parachute compartments Inspection 2.11 The launch vehicle must have an electronic tracking device 2.12 Recovery system electronics shall not be adversely affected by any other on- board electronic device Altimeter compartments are shielded by intenal copper mesh lining Inspection, Testing 2.12.1 Altimeters for the recovery system must be in a separate compartment than any other transmitting device Each altimeter bay has a dedicated and separate compartment Inspection
  39. 39. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 39 2.12.2 Recovery electronics shall be shielded from all on-board transmitting devices Altimeter compartments are shielded by intenal copper mesh lining Inspection,Testing 2.12.3 Recovery electronics shall be shielded from any magnetic waves generated by on- board devices Altimeter compartments are shielded by intenal copper mesh lining Inspection, Testing 2.12.4 Recovery electronics shall be shielded from any on-board device that could adversely affect proper operation Altimeter compartments are shielded by intenal copper mesh lining; Each altimeter bay has a dedicated and separate compartment Inspection,Testing 2.13 Recovery system shall use commercially available low- current e matches for ignition of ejection charges Davyfire N28BR e- matches have been selected Inspection, Testing Table 10 Recovery System Verification Table Risks and Plans for Reducing Risks System Risks: Each system has specialized risks. Certain risks are more likely to occur than others, and some risks have a more severe consequence. In order to avoid the realization of a risk, mitigation is performed. Table 11 lists some of the specific risks to each system, the risk’s likelihood, severity, consequence, and mitigation. System Risk Likelihood Severity Consequence Mitigation Safety Unable to Obtain Flight Waivers Medium Medium Delay in Test Launches Schedule Test Flight at Official Test Launch Sites Disregarding Safety Plan Low High Harm to Participants Strict Safety Plan
  40. 40. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 40 Enforcement Ignorance of Legal Rules and Regulations Low High Legal Repercussions Recurring Legal Education Structure Destruction of Parts During Testing High Low Reordering of Parts, Delay in Project Inventory Extra Parts Improper Test Vehicle Assembly Medium Medium Unpredictable Performance Assembly Checklist Propulsion Inconsistent Motor Construction Low High Unpredictable Performance Thorough Inspection and Analysis Incorrect Motor Mount Assembly Low High Possible Harm to Vehicle and Participants Assembly Checklist Educational Engagement Poor Quality Presentation Low High Less Effective Outreach Rehearsed Presentation Scheduling Conflict With The Schools Medium High Cancellation of Events Communicate With Schools Electronics Faulty Components Low Medium Delay to Project, Ineffective Payload Ordering Duplicates, Testing Incorrect Wiring Configurations Low Low Harm to Parts, Ineffective Payload Thorough Research and Design Power Budget Miscalculation Low High Harm to Parts, Ineffective Payload Redundant Calculations Recovery Incorrect Black Powder Rating Low High Harm to Vehicle and Participants Extremely Thorough Testing Parachute Size Miscalculation Low Low Violation of Competition Rules Redundant Calculations Public Relations Lack of Project Exposure Low Medium Lack of Funding and Support Constant Outflow of Updates and Information Harmful Representation of Team Low Medium Lack of Funding and Support Proper Ethics and Professionalis m Management Inadequate High Medium Delay in Weekly Team
  41. 41. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 41 Communicatio n Project, Project Failure, Design Flaws Meetings, Email Minutes Scheduling Conflicts Medium Medium Delay in Project Maintain Scheduling University Administration Delays Medium Low Delay in Project/Fundin g Proper Communicatio n and Scheduling Software Loss of Source Code Medium High Rewrite Software Backup Source Code Incorrect Algorithm Design Low Medium Ineffective Payload Logic Redundant Calculations Incorrect Datasheets Low Low Confusion, Delay in Project Datasheet Verification, Testing Table 11 System Risks Project Risks: There are many risks to the overall project. Each risk has a certain level of likelihood and severity. The consequences of such risks effect the overall operation of the team. It is important to mitigate such risks and decrease their likelihood. Figure 12 displays the risk plot, where all risks in consideration above plot according to the likelihood they will occur versus the severity of their consequence. Table 12 lists each risk with an associated number on the plot such that every risk is identifiable on the plot.
  42. 42. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 42 Severity # Severity Probabilty Severity # Severity Probabilty Severity # Severity Probability LOW Medium High 20 Delay in Deliveries 90 60 Communication Failure 10 90 InadequatePersonnel 25 33 Destruction of Parts During Testing 85 66 Unableto Obtain FlightWaivers 0 80 Manufacturing Issues 20 1 IncorrectWiring Configurations 50 45 Improper TestVehicleAssembly 15 99 NAR/TRAViolations 5 3 ParachuteSizeMiscalculation 10 39 Faulty Components 4 98 Damageof Property 15 30 University Administration Delays 40 33 Lack of ProjectExposure 20 70 OSHAViolations 10 5 IncorrectDatasheets 5 34 Harmful Representation of Team 10 100 Personal Injury 1 50 InadequateCommunication 66 67 Teammates Disregarding Safety Plan 2 55 Scheduling Conflicts 75 99 Ignoranceof Legal Rules and Regulations 3 35 IncorrectAlgorithmDesign 55 72 InconsistentMotor Construction 22 57 EnvironmentPrevents Recovery 25 75 Loss of SourceCode 50 52 Poor Quality Presentation 10 33 Poor Weather 18.5 29 Scheduling ConflictWith TheSchools 12 66 Burn Ban in Effect 25 42 Power BudgetMiscalculation 1 92 Loss of RocketLab 5 93 Loss of Low-AltitudeTestLaunch Facility in Glen Rose,TX 5 94 Loss of High-AltitudeTestLaunch Facility in Cross Plains,TX 5 38 Loss of ScienceBuilding 0 39 Engineering Building 0 40 Loss of HCI facilities 12 Figure 12 - Risk Plot Table 12 Project Risks
  43. 43. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 43 Planning of Manufacturing The team has two Industrial Technology majors, experienced in detailed CAD design including 3D modeling as well as manufacturing/machining with CNC machines. The team uses the university’s manufacturing facilities to mill, drill, cut, lathe or machine parts. Members have proper training in the safe operation of such machines. Access to this facility will allow the team to make adjustments to the structure design. This is a time efficient alternative to ordering manufactured structure components. Planning of Verification The verification table (Table 9, pages 32-36) indicates all respective requirements processes for verification. This table guides the group throughout all testing phases, and all test data analysis ensures that all requirements are in a state of satisfaction. The Testing Summary Table (Table 13) illustrates a summary of the testing that is pending. For dates and deadlines of verification testing, refer to the project plan section with the testing Gantt chart, Figure 60. Testing Title Subsystem Structure Testing Structure Lab Prototyping Structure, Recovery, Integration Low Altitude flight Structure, Recovery Dual deployment Recovery Force of impact Structure, Recovery, Integration Full scale launch Vehicle, Recovery, Integration Timed final assembly Structure, Recovery, Integration Table 13 Testing Summary Structure testing is at Harris Composites’ testing facilities in Granbury, TX. This company performs materials strength testing. Arrangements exist to test the materials that make up the vehicle’s airframe. In particular, they provide data on material strength and integrity. These figures verify vehicle performance under expected loads. Lab prototyping takes place on all subsystems of the vehicle to ensure compatibility and feasibility, as well as to identify any immediate flaws in the design or manufacturing. This includes bench top testing, representative model and prototype builds, as well as documenting and modifying the changes to the design as needed. Low altitude flights take place for proof of concept performance, where critical
  44. 44. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 44 functionality of the vehicle can be verified. Motors for this test are less expensive than a full scale launch, and flight waivers are more readily available. This allows adequate test flights to take place and sufficient test data to generate for diagnostics. The relative ease in conducting this low altitude flights makes this a very valuable test mode. Dual deployment testing takes place to ensure functionality of the recovery system. This includes testing on the ground to verify separation events and parachute ejection, as well as perfecting ejection charge specifications. Dual deployment is in the plan for all test flights, both low altitude and full scale. Force of impact testing will take place to analyze and verify the structural integrity of the vehicle at landing, as well as the functionality of the recovery system. The team employs a testing accelerometer in drop testing, as well as in test flights. This will provide a quantitative justification that the recovery system is sufficient to meet all requirements and the structural design and integration of the vehicle is adequate. Full scale test launches takes place to verify overall functionality of the vehicle. The actualization of these is to represent the competition conditions, and one of these in particular is useful as the full scale demonstration flight before the FRR. This test is extremely vital to confirming the design, as it requires all respective components of the vehicle to perform as intended. Timed assemblies of the launch vehicle ensure that all components integrate and that the vehicle is ready for a launch within the time limit at competition. Tuning up the team for timely assemblies reduces the chance for potential error in preparing the vehicle for launch. Planning of Integration The project manager and lead engineer are present at all team meetings and subsystem meetings. They carry the responsibility of ensuring proper and efficient communication of the group, such that all necessary subsystems of the vehicle design integrate successfully according to the plan. Three team meetings per week are part of the general schedule, which allows the subsystem leads to present their progress in front of the entire team. These sessions allow the team to address any design issues, concerns, or questions. Planning of Operations All operations relating to this project must have a schedule, procedure, and checklist to ensure all steps leading to the successful completion of each operation happens. This ensures efficiency in carrying out a particular operation and allows checklists and procedures to develop in accordance with safety regulations. All
  45. 45. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 45 test flights follow the checklist found in Appendix B. Prior to each flight, testing verification on each component takes place to ensure flight readiness. The team is divided into several subsystems. Figure 13 is a hierarchy chart of the team. Figure 13 - Team Hierarchy Confidence and Maturity of Design The vehicle design evolves throughout the design process. Every change seeks to improve the overall design. Over time, the flaws and failures of the design lose out, helping the design to mature. The more time analyzing and testing should increase the opportunities for evolution. The overall maturity of the design should increase throughout these evolutionary stages. Already, the design has progressed through several evolutionary phases. The current design is at an intermediate maturity. While inevitably many design flaws persist, many cease to be and changes continually mount to improve upon the design. It is necessary to follow through on planning and execution of necessary maturity/risk reduction efforts throughout the product life cycle. Figure 14 shows the maturity life cycle of the project.
  46. 46. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 46 Figure 14 - Project Life Cycle The Tarleton Aeronautical Team is confident in their design abilities. Although it is certain the design will evolve, the current design has progress to speak for itself. Extensive hours formulating the current design have been worth the effort. The team is confident in the overall design of the vehicle and the ability to achieve the target goals in the competition. The design chosen for this year’s competition is simple and efficient, with a clear, modular payload system. The team examines every subsection in detail for flaws or possible improvements on a theoretical level. Tests on each component continually reveal further information regarding the design’s maturity. Dimensional Drawing The major sections of the vehicle are represented in Figure 15. Figure 15 - Dimensional Drawings
  47. 47. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 47 Electrical Schematics of the Recovery System The recovery system consists of three main electrical devices: the PerfectFlite Stratologger recovery altimeter, the Featherweight Raven 3 recovery altimeter, and the BeeLine GPS. The recovery altimeters each utilize a dedicated nine volt power supply, and the BeeLine GPS utilizes a five volt power supply. Correct wiring of the recovery altimeters is crucial to a safe and successful recovery. Improper wiring could cause inadvertent deployment of the recovery system, risking injury to people and the vehicle. Figure 16 is a conceptual wiring schematic of the recovery’s electronic components. There are two setups of the recovery system electronics; one for the drogue deployment and one for the main. Furthermore, there are two recovery GPS modules. Figure 16 - Recovery Electrical Schematic Figure 17 is a picture of the Featherweight Raven 3 deployment altimeter.
  48. 48. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 48 Figure 17 - Featherweight Raven 3 Wiring Mass Statement The mass summary of the vehicle is in Table 14. Each subsection breaks down into its respective components in Tables 15 through 17. The mass calculations for the launch vehicle, subsections, and individual components come from three methods. The mass of components is retrievable from data sheets when available. Density of the materials and volume of the structural components help obtain mass estimates. Where no data is available, logical deductions provide reason towards the component mass based on similarity to other known components. This allows for a reasonable level of accuracy and to allow a reserve of three to five pounds for a possible mass growth. Concluding from the listed mass of 33.5 lbs for the launch vehicle and the maximum thrust of 437.7 lbf from the propulsion system, the rocket has a thrust to weight ratio of 13:1. This requires more than 400 lb of additional mass to prevent the vehicle from launching. Overall Subsection Mass (oz) Mass (lb) Payload 40.58 2.54 Recovery 59.85 3.74 Structure 435.6 27.23 Total Mass (Launch) 536.03 33.50 Total Mass (Apogee) 473.92 29.62 Table 14 Total Mass Summary
  49. 49. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 49 Payload Component Quantity Mass (oz) Total Mass (oz) Baseplate 1 1.11 1.11 Battery 8 1.28 10.24 Circuit Boards 1 6 6 Railing – Main 2 2.8125 5.625 Railing – Support 2 0.262 0.524 Sensors/Electronics 1 13.1 13.1 Servo – Large 1 1.55 1.55 Servo – Small 1 0.67 0.67 Video Camera 1 1.76 1.76 Subtotal 40.579 Table 15 Payload Mass Summary Recovery Component Quantity Mass (oz) Total Mass (oz) Attachment Hardware 2 3 6 Charges – Drogue 1 3 3 Charges – Main 1 4 4 Deployment Bag – Drogue 1 3 3 Deployment Bag – Main 1 5 5 GPS 2 2 4 Parachute – Drogue 1 2.63 2.63 Parachute – Main 1 11.3 11.3 Recovery Electronics – Drogue 1 5 5 Recovery Electronics – Main 1 5 5 Shock Cord – Drogue 1 4.68 4.68 Shock Cord – Main 1 6.24 6.24 Subtotal 59.85 Table 16 Recovery Mass Summary Structure
  50. 50. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 50 Component Quantity Mass (oz) Total Mass (oz) Acrylic Payload Section 1 52.3 52.3 Ballast 1 10.92 10.9 Bulkhead 3 3.03 9.09 Bulkhead – Motor 1 6.07 6.07 Bulkhead – Payload 2 32.8 65.6 Center Rings 3 2.01 6.03 Coupler 2 14.3 28.6 Engine Compartment 1 12.9 12.9 Body Tube – Front 1 38.4 38.4 Body Tube – Rear 1 49.4 49.4 Fin 4 5.625 22.5 Motor 1 118 118 Nosecone 1 15.8 15.8 Subtotal 435.6 Table 17 Structure Mass Summary Recovery System Deployment of Parachutes A dual-stage deployment recovery system is in use, consisting of the staged release of a drogue parachute and a main parachute. The main parachute ejects from the top of the upper body structure, just below the nose cone. The drogue parachute ejects from the drogue parachute compartment at the front of the lower body structure. This staging is in Figure 18.
  51. 51. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 51 In order to minimize both landing radius and terminal velocity, the drogue parachute deploys at 5,280 feet, when the vehicle is at apogee. The main parachute deploys at 500 feet above ground level on descent to ensure proper final velocity. It is imperative that the drogue parachute deploys at apogee in order to avoid damage to the rocket body caused by the jarring that would ensue due to a high speed ejection. Deployment Altimeters In order to eliminate the variability of choosing the right delay time and to improve redundancy, each deployment functions with two altimeters. Each altimeter system consists of a main altimeter, backup altimeter, and e-match wiring. The main altimeter is a Featherweight Raven3 and the backup is a PerfectFlite StratoLogger which is completely independent of the payload electronics. Each altimeter system is inside of a vented compartment below each parachute compartment in the vehicle body. Each altimeter has its own dedicated power supply, a standard 9-volt battery. Each altimeter system mounts vertically on a 0.125-inch-thick, 4-inch-wide, 1-inch-long fiberglass board. One altimeter is on each side. Each board then epoxies on either end to a 0.125-inch-thick, 5.375- inch-diameter fiberglass disk. The entire setup bolts to the bulkhead below each parachute compartment. The compartments seal from the black powder ejection charges. Each compartment vents to ambient air pressure in order to acquire proper altitude readings. There is a porthole drilled from the exterior of the rocket body into each altimeter compartment, the size of which shall be determined through later testing. Item Distributor Product Unit Dimensions Unit Cost Number Total Cost Main Altimeter Featherweight Altimeters Raven3 1.8in. X 0.8in. X 0.55in. X 0.34oz $155.00 2 $310.00 Loki Research Ozark ARTS 3.75in. long X 1.4in. wide X 2.75 oz $190.00 2 $380.00 Backup Altimeter PerfectFlite Stratologger 2.75in. long X 0.9 in. wide X 0.45 oz $79.95 2 $159.90 Adept Rocketry ALTS1-50K 0.9in. X 0.65in. X 4.25in. X 4.25oz $89.00 2 $178.00 Table 18 Deployment Altimeter Trade and Selection
  52. 52. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 52 The Featherweight altimeter is the main altimeter due to its versatility. It has full functionality regardless of positioning, has visible and audible readout of individual channel continuity and battery voltage, allows for user calibration of the accelerometer rather than presets, can record up to eight minutes of high-rate data plus an additional 45 minutes per flight, and has a downloadable interface program which is easy to read. The audible readout function deactivates manually prior to launch. Table 18 shows that the Featherweight is the choice for the project. Though features are comparable, physical dimensions are not. Optimal engineering efficiency comes through employment of the Featherweight. The cost of the Featherweight is less than that of the Ozark. The PerfectFlite altimeter is the backup altimeter due to its high level of reliability. False triggering is not a problem for gusts of wind up to 100 miles per hour. The precision sensor and 24-bit analog-to-digital converter (ADC) allow for 99.9 percent accurate altitude readings, and the selectable apogee delay for dual setups prevents overpressure from simultaneous charge firing. As demonstrated in Table 18, the PerfectFlite proves superior to the ALTS1-50K. While features compare well, the ALTS1-50K is not the best choice concerning compartment space capacity or cost. Additionally, each altimeter system has an externally-accessible magnetic arming switch capable of being locked in the “on” position for launch. The arming switch dedicated to the dual altimeters, which control main parachute deployment, are at five feet, eight inches above the base of the launch vehicle. Those dedicated to the dual altimeters which control drogue parachute deployment are two feet above the base of the launch vehicle. Ejection Charges In order to ensure separation and ejection of the proper parachute at the proper time, each altimeter is set to light a one-foot low-current Daveyfire N28BR electric match. There are holes between the lower bulkhead and altimeter compartment to allow the lead on each e-match through; the holes are shut with epoxy to seal the chamber from the other chambers in the vehicle body. The basic construction of Figure 19 Daveyfire Electric Match
  53. 53. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 53 an electric match, or e-match, is in Figure 19. The e-match selection is indicated in Table 19. While the Daveyfire is more costly than the QuickBurst or RocketFlite, it is the only fully-assembled option. The QuickBurst and RocketFlite are less costly options, but each requires manual assembly. Improper assembly of an e-match could result in electric shock and premature ignition of ejection charges. Distributor Item Unit Dimensions Pre- Assembled Unit Cost Coast Rocketry Daveyfire N28BR 1 ft long X 1 Yes $2.95 QuickBurst QuickBurst E-Match Kit 1 ft long X 20 No $32.00 RocketFlite MF-12 1 ft long X 12 No $9.95 Table 19 Electric Match Trade & Selection Assuming that the entire mass of each charge is burns and converts into a gas, the basic Ideal Gas Law is used, ( ) With , and for the volume of the cylinder, ( ) ( ) ( ) ( ) ( ) For a compartment length of for the drogue ejection,
  54. 54. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 54 For a compartment length of for the main ejection, Each black powder ejection charge consists of a well of black powder contained in a plastic charge tube with the shroud of an e-match immersed in the well. Global Positioning System A BeeLine GPS is the selection for recovering each component upon landing in the event that tethering separation on descent occurs or visual contact is lost. Each GPS is in a 2.5 inch sub-compartment of each parachute compartment. In order to achieve this, each GPS mounts to a 0.125 inch thick, 2.25 inch long, four inch wide sheet of fiberglass with epoxy, then on either end to a 0.125 inch thick and 5.375 inch wide fiberglass disk which is inserted below the lower bulkhead of each altimeter compartment. This should shield the devices from parachute ejection, black powder charge ignition, and fuel ejection. A fine copper wire mesh lines the internal surface of the drogue altimeter housing, which is separate from the GPS sub-compartment via bulkhead, such that these altimeters are shielded from radio frequencies in order to prevent inadvertent excitation. The holes in this mesh must be significantly smaller than the wavelength of the interfering radio frequencies so that the enclosure does not ineffectively approximate an unbroken conducting surface. The BeeLine GPS operates in the range 420 to 450 mHz. The XBee operates at 900 mHz. A pure copper mesh fabric with electromagnetic frequency blocking effectiveness in the range 900 to 420 mHz is the team’s choice to line each compartment containing a BeeLine GPS. A corresponding ground receiver is in the ground station. Each BeeLine package includes a fully integrated RF transmitter, GPS and RF antennas, GPS Module, and battery. Altogether, these devices simultaneously transmit latitude, longitude, altitude, course, and speed. These quantities are analyzable after each flight in order to aid in continued optimization. The BeeLine GPS has been chosen for its small size, reasonable cost, transmission range at up to 20 miles line of sight, frequent usage in high-powered model rocketry, use of standard decoding hardware (automatic packet reporting system, or APRS), and operation frequency on any frequency in the 70-centimeter amateur radio band. Additionally, the BeeLine is the only consistently commercially available fully integrated GPS system for model rockets. Its measurements of course and speed allow for real-time calculation of landing distance and terminal kinetic energy. These measurements also serve as a check for the altimeters. The mounting precautions, the 8 hour battery life of the Lithium-Poly battery, the non- volatile flight-data memory storage (3 hours at 1 Hertz), the user-programmable
  55. 55. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 55 transmission rates, and output power will ensure that the devices remain fully functional during the course of the flight. Recovery Testing Static testing on each set of altimeters with e-matches ensures the reliability of the electronic system. The possibility of delaying the signal from the backup altimeter by up to two seconds with respect to the main altimeter is under consideration. This could help to ensure that separation does occur should black powder well leak, humidity become a problem, or pressure conditions prevent a sufficiently powerful force from the main charge. Once static testing on the altimeter-e-match systems is complete, ground testing of the system with the charge wells in an empty replica of the launch vehicle is conducted under the guidance of the team mentor in the presence of the team safety officer. This enables any sizing adjustments which may be necessary to ensure vehicle separation are addressed prior to conducting a full-scale launch. The final testing phase of the recovery system components includes at least one subscale single-event flight, at least one subscale dual-event flight, and lastly at least one full-scale dual-event flight. The purpose of the subscale single-event flight is to allow for proper understanding of the parachute components in combination with the electronic charge system. Once the parachute components perform in real a situation, a subscale dual-event flight will serve to mimic the competition flight in a more contained way. Finally, a series of full-scale dual-event flights confirms expectations for the competition flight. The cycle of these tests is in Figure 62, but specific data to the recovery system is in Table 20, where one denotes static testing, two denotes sub-scale, and three denotes full-scale. Figure - Recovery Testing Dates Month Date Stage Verification Event October 27 2 Dual Deployment Test Launch November 12 1 Parts Ordered for Prototyping 17 3 Dual Deployment Launch 30 1 End of Lab Prototyping December 1 2 Low-Altitude Flight 3 3 Motor Assembly 5 1 End of Programming 8 3 Full Scale Launch 22 2 Low Altitude Flight January 2 2 End of Field Testing 5 3 Test Launch
  56. 56. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 56 12 3 Alternative Launch 19 1 Static Motor Test 26 3 Low Altitude Full Force of Impact Test Launch Table 20 Recovery Testing Dates Anemometer In order to anticipate the effectiveness of the recovery system prior to flight, a compact rotary-fan digital anemometer is incorporated into the ground station. As indicated in the below Table 21, the SpeedTech WindMate-300 (WM-300) is the team’s choice for its low price in comparison to newer, similar products. Like the more costly options listed, the WM-300 provides a digital readout of wind speed, wind direction, humidity, and pressure. It is also resistant to water damage and has a threaded base which may be mounted to a tripod at the ground station. Mounting the anemometer to a tripod maintains stability of the device to ensure accurate readings. Component Distributor Item Unit Dimensions Wind Speed and Direction Water Proof Mount for Tripod Unit Cost 4500 Pocket Weather Tracker Ambient Weather Kestrel 4500 5in. X 1.8in. X 1.1in. Yes Yes Yes $299.00 WindMate Anemometer Weather Shack Speed Tech WM- 350 5.5in. X 1.75in. X 0.75in. Yes Yes Yes $229.95 WindMate Anemometer Weather Shack Speed Tech WM- 300 5.5in. X 1.75in. X 0.75in. Yes Yes Yes $154.95 Table 21 Anemometer Trade & Selection Parachute Size Calculations Assuming use of a Cesaroni L1720-WT-P motor, the total vehicle launch weight is estimated to be 29.62 pounds, as seen in Table 14. The team takes into consideration the addition of up to 10 percent ballast, as well as the fuel compartment being empty by deployment of the drogue parachute, and estimates the vehicle weight at 29.62 pounds. While the weight estimate may change, the process of the following calculations will not.
  57. 57. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 57 The following calculations give the maximum descent rate upon landing for the vehicle to have a kinetic energy of less than 75 foot pound force. ( ) √ ( ) Therefore for the vehicle to land with a kinetic energy of less than 75 foot pound force, based on a weight of 29.18 pounds, the descent rate must be less than 12.8656 feet per second. Using this result, it is possible to calculate the minimum diameter of the main parachute as follows: ( ) ( )
  58. 58. Tarleton Aeronautical Team 2012 - 2013 USLI Preliminary Design Review 58 ∑ . ( ) ( ) ( ) √ √ Therefore, for the vehicle to land with a kinetic energy of less than 75 foot pound force, the main parachute must be at least 9.965 feet in diameter. Due to commercial availability, the main parachute diameter is 10 feet. The following calculates the exact descent rate of the vehicle based on the 10 foot diameter main parachute.

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