Tarleton Aeronautical Team i 2012 - 2013 USLI Preliminary Design Review

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.

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Tarleton Aeronautical Team ii 2012 - 2013 USLI Preliminary Design Review

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

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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

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Tarleton Aeronautical Team iv 2012 - 2013 USLI Preliminary Design Review

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

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Participation Goal ...... 150 Accomplished Educational Outreach ...... 150 Conclusion ...... 153 Appendix A - Itemized Subsystem Budget ...... 155

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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

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Tarleton Aeronautical Team vii 2012 - 2013 USLI Preliminary Design Review

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

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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

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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.

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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

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Tarleton Aeronautical Team xii 2012 - 2013 USLI Preliminary Design Review

TAP-Technical Advisor Panel TRA-Tripoli Rocketry Association TTL Transistor-transistor logic USLI-University Student Launch Initiative UV-Ultraviolet

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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 Primary Featherweight Raven 3 Altimeter, Avionics 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 Gather Atmospheric and GPS Data, Autonomously Orientate Photographic Directorate Payload Camera, Capture Video for Public Outreach, and a Clear Acrylic Housing

Table 2 Payload Summary

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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

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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.

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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 Payload section integration and design is better suited for the 5.525 inches wider diameter

To allow some additional room for recovery system Total length from 103 in to 108 in implementation

Fin size and design To maintain stability for a larger vehicle

Incorporation of a ballast system in the nose To allow a greater level of ballast weight control on launch cone 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 To provide shielding from wireless transmitting devices compartments 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

To provide more room for autonomous camera Payload width from 3.5” to 5" orientation system and inclusion of a video camera

Payload housing will be mounted by screws Avoids manufacturing a threaded acrylic cap through the acrylic to a fiberglass coupler

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Inclusion of servos and an accelerometer for the Critical payload objective autonomous camera orientation system

To reduce overall space required as compared to PCB planning and design perforated board mounting

The official scoring altimeter moved from recovery To avert pressure fluctuations from ejection charges, to bays to payload housing 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 For structural support milled grove of the upper payload mounting cap Two, 90 degree opposing boards for solar Solar irradiance and UV sensors can gather data at 90 irradiance and UV sensors degree intervals around the rocket vehicle Handheld Yagi directional antenna will be used to Development of an automated tracking system for the receive all transmitted data from the payload 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 To increase quality of Educational outreach Educational Outreach Plan has been further developed aspect to the entire project To increase quality of public outreach and Community Outreach events added project sustainability Table 5 Project Plan Changes

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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

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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

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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.

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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.

Peak Load Peak Stress Avg. Cost per Availability Material Modulus (lbf) (psi) Foot (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

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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 Separate Fiberglass Extended Upper Acrylic Options Airframe Airframe Airframe

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Visibility of electronics Durability, cost, and payload Durability, cost, availability, familiarity, Pros components from availability, familiarity lack of coupling system exterior, internally (lighter) mounted sensors

Inability of external TableInability 7 Pros of and external Cons of Payload Options inspection, externally inspection, externally Cons mounted sensors, Durability, weight mounted sensors, inaccessibility to coupling system internal systems

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

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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.

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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:

Selected Option Alternative Option 1 Alternative Option 2

Figure 7 - Fin Design Options

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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 Figure 8 Selected Fin Design

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

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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.

Thrust Velocity Max. Burn Availability Apoge Total Average to Motor Off Rail Velocity Time / e (ft.) Impulse Thrust Weight (ft./s) (ft./s) (s) Cost Ratio 830.9lbf Cesaroni s 394.3lbf High/ L1720- 5345 75.9 738 2.15 11.8 (3696Ns (1754N) $170.96 WT-P ) 887.8lbf Aerotech s 308.9lbf Medium/ 5618 65.6 723 2.65 9.2 L1390G (3949Ns (1374N) $209.99 ) Cesaroni 1082lbfs 246.6lbf Medium/ L1090SS- 6479 76.8 (4815Ns 733 4.4 7.4 (1097N) $346.95 P ) 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.

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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

. ( ) ( )

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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.

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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

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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.

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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.

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Requirement Satisfying Design Verification Vehicle Requirement (SOW) Feature Method

Vehicle shall deliver payload to Testing, 1.1 Motor selection 5,280 feet AGL Analysis

Altimeter Model X is used Vehicle shall carry one official 1.2 and included in the SMD Inspection scoring barometric altimeter payload section

Official scoring altimeter shall Altimeter Model X has Testing, 1.2.1 report the official competition this functionality Inspection altitude via a series of beeps

Four additional altimeters, outside of the Teams may have additional 1.2.2 payload, will be used to Inspection altimeters detect apogee and ignite ejection charges

At Launch Readiness Review, a Official altimeter 1.2.2.1 NASA official will mark the placement allows ease of Inspection altimeter to be used for scoring locating and marking

At launch field, a NASA official The official altimeter has Testing, 1.2.2.2 will obtain altitude by listening to this functionality Inspection beeps reported by altimeter

At launch field, all audible No other electronics in electronics except for scoring 1.2.2.3 the design have audible Inspection altimeter shall be capable to turn indicators off

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Official, marked altimeter is Functional Recovery Testing, 1.2.3.1 damaged and/or does not report System Inspection an altitude with a series of beeps

Team does not report to NASA This task will be assigned official designated to record Analysis, 1.2.3.2 to an appropriate team altitude with official marked Inspection member altimeter on launch day

Altimeter reports apogee altitude Testing, 1.2.3.3 Motor selection of over 5,600 feet Analysis

Launch vehicle remains subsonic Testing, 1.3 Motor selection from launch until landing Analysis

Testing, Vehicle must be recoverable and Recovery system allows 1.4 Inspection, reusable a safe landing of vehicle Analysis

Launch vehicle shall have a Vehicle is composed of 3 1.5 maximum of four independent Inspection tethered sections sections

Launch vehicle shall be prepared Launch operations and Testing, 1.6 for flight at launch site within 2 assembly procedure Inspection hours

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All critical on-board Launch vehicle will remain components will have Testing, 1.7 launch-ready for a minimum of sufficient capacity to Inspection, one hour with critical functionality meet entire system Analysis runtime (1.5 hours)

Vehicle shall be compatible with 1010 rail buttons 1.8 either 8 feet long 1 inch rail Inspection attached to vehicle body (1010)

Launch vehicle will be launched Motor selection is 1.9 by a standard 12 volt DC firing compatible with this firing Inspection system system

Launch vehicle shall require no Motor ignition only 1.10 external circuitry or special requires the 12V DC Inspection equipment to initiate launch firing system

Launch vehicle shall use a 1.11 commercially available, certified Cesaroni L1720 Inspection APCP motor

Total impulse provided by launch 1.12 vehicle will not exceed 5,120 3695.6 Ns Inspection Newton-seconds

The full scale rocket, in final flight configuration, must be 1.15 Testing Schedule Testing successfully launched and recovered prior to FRR Featherweight and Vehicle and recovery system 1.15.1 Stratologger Altimeters, Inspection function as intended Parachute Calculations

The schedule allows for Payload does not have to be 1.15.2 the payload to be flown in Schedule flown during full-scale test flight. the full scale launch.

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If payload is not flown, mass If this occurs mass will be 1.15.2.1 simulators shall be used to Testing added proper sections. simulate payload mass

Mass of each section is Mass simulators shall be located calculated, and mass will 1.15.2.1.1 in same location on rocket as the Testing be added in the missing payload mass appropriate sections.

Any energy management system There will be no changes or external changes to the 1.15.2.2 to the external surface of Design surface of the rocket shall be the rocket. active in full scale flight

Unmanned aerial vehicles, and/or recovery systems that control 1.15.2.3 flight path of vehicle, will fly as N/A N/A designed during full scale demonstration flight Full scale motor does not have to The schedule and budget 1.15.3 be flown during full scale test plan for the full scale Schedule flight motor to be flown.

Vehicle shall be flown in fully The schedule plan for the 1.15.4 ballasted configuration during full Schedule fully ballasted system. scale test flight

Success of full scale demonstration flight shall be Pat Gordezlick will be documented on flight certification 1.15.5 present at the full scale Schedule form, by a Level 2 or Level 3 launch. NAR/TRA observer, and documented in FRR package

After successfully completing full- scale demonstration flight, launch The schedule plans for vehicle or any components shall the design to be 1.15.6 Schedule not be modified without complete and changes to concurrence of the NASA Range be complete. Safety Officer (RSO) Budget indicates that the Maximum amount teams may total spent on the rocket Inspection, 1.16 spend on rocket and payload is and payload is less than Analysis $5000 $5000.

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Table 9 Vehicle Verification Table

Table 10 is the verification table for the recovery system.

Requirement Satisfying Design Verification Recovery System (SOW) Feature Method Requirement

Recovery devices shall be Altimeters will stage staged such that a drogue ejection charges for 2.1 parachute is deployed at Testing respective parachutes at apogee; main parachute is at the prescribed altitudes 500ft.

Each independent section of Main parachute 2.2 the launch vehicle will have a Testing, Analysis selection maximum KE of 75 ft-lbf.

Each independent section of Drogue and main 2.3 the vehicle shall land with Testing, Analysis parachute selection 2500 ft. of the launch pad

Recovery circuits are Recovery electrical circuits independent of payload 2.4 shall be independent of Inspection with dedicated power payload electronics supplies

Each deployment event Recovery system must 2.5 is controlled by a main Inspection include redundant altimeters and backup altimeter

Each altimeter shall be Port holes in vehicle 2.6 armed in launch configuration airframe to altimeter Inspection with external arming switches bays

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Each altimeter shall have a Each altimeter uses a 2.7 Inspection dedicated power supply separate 9 Volt battery

Port holes in vehicle Each arming switch can be Testing, 2.8 airframe to altimeter locked in the ON position Inspection bays

Main altimeter bay is located at 6 feet above Each arming switch will be a the base; drogue 2.9 max of 6ft above the vehicle Inspection altimeter bay is located base at 2 feet 4.5 inches from the base

Removable shear pins shall Nylon shear pins will be be used for the drogue and used to couple 2. 10 Inspection main parachutes parachute compartments compartments

The launch vehicle must 2.11 have an electronic tracking device

Recovery system electronics Altimeter compartments shall not be adversely Inspection, 2.12 are shielded by intenal affected by any other on- Testing copper mesh lining board electronic device

Altimeters for the recovery Each altimeter bay has system must be in a separate 2.12.1 a dedicated and Inspection compartment than any other separate compartment transmitting device

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Recovery electronics shall be Altimeter compartments 2.12.2 shielded from all on-board are shielded by intenal Inspection,Testing transmitting devices copper mesh lining

Recovery electronics shall be Altimeter compartments shielded from any magnetic Inspection, 2.12.3 are shielded by intenal waves generated by on- Testing copper mesh lining board devices

Altimeter compartments Recovery electronics shall be are shielded by intenal shielded from any on-board copper mesh lining; 2.12.4 Inspection,Testing device that could adversely Each altimeter bay has affect proper operation a dedicated and separate compartment

Recovery system shall use Davyfire N28BR e- commercially available low- Inspection, 2.13 matches have been current e matches for ignition Testing selected of ejection charges

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 Schedule Test Unable to Flight at Obtain Flight Delay in Test Official Test Safety Waivers Medium Medium Launches Launch Sites Disregarding Harm to Strict Safety Safety Plan Low High Participants Plan

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Enforcement Ignorance of Legal Rules and Legal Recurring Legal Regulations Low High Repercussions Education Destruction of Reordering of Parts During Parts, Delay in Inventory Extra Testing High Low Project Parts Structure Improper Test Vehicle Unpredictable Assembly Assembly Medium Medium Performance Checklist Inconsistent Thorough Motor Unpredictable Inspection and Construction Low High Performance Analysis Propulsion Incorrect Possible Harm Motor Mount to Vehicle and Assembly Assembly Low High Participants Checklist Educational Poor Quality Less Effective Rehearsed Engagement Presentation Low High Outreach Presentation Scheduling Conflict With Cancellation of Communicate The Schools Medium High Events With Schools Delay to Project, Ordering Faulty Ineffective Duplicates, Electronics Components Low Medium Payload Testing Incorrect Harm to Parts, Thorough Wiring Ineffective Research and Configurations Low Low Payload Design Harm to Parts, Power Budget Ineffective Redundant Miscalculation Low High Payload Calculations Harm to Extremely Incorrect Black Vehicle and Thorough Recovery Powder Rating Low High Participants Testing Violation of Parachute Size Competition Redundant Miscalculation Low Low Rules Calculations Constant Lack of Outflow of Public Lack of Project Funding and Updates and Relations Exposure Low Medium Support Information Proper Ethics Harmful Lack of and Representation Funding and Professionalis of Team Low Medium Support m Management Inadequate High Medium Delay in Weekly Team

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Communicatio Project, Project Meetings, n Failure, Design Email Minutes Flaws Scheduling Delay in Maintain Conflicts Medium Medium Project Scheduling Proper University Delay in Communicatio Administration Project/Fundin n and Delays Medium Low g Scheduling Loss of Source Rewrite Backup Source Software Code Medium High Software Code Incorrect Algorithm Ineffective Redundant Design Low Medium Payload Logic Calculations Confusion, Datasheet Incorrect Delay in Verification, Datasheets Low Low Project 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.

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Figure 12 - Risk Plot

Severity # Severity Probabilty Severity # Severity Probabilty Severity # Severity Probability LOW Medium High 20 Delay in Deliveries 90 60 Communication Failure 10 90 Inadequate Personnel 25 33 Destruction of Parts During Testing 85 66 Unable to Obtain Flight Waivers 0 80 Manufacturing Issues 20 1 Incorrect Wiring Configurations 50 45 Improper Test Vehicle Assembly 15 99 NAR/TRA Violations 5 3 Parachute Size Miscalculation 10 39 Faulty Components 4 98 Damage of Property 15 30 University Administration Delays 40 33 Lack of Project Exposure 20 70 OSHA Violations 10 5 Incorrect Datasheets 5 34 Harmful Representation of Team 10 100 Personal Injury 1 50 Inadequate Communication 66 67 Teammates Disregarding Safety Plan 2 Ignorance of Legal Rules and 55 Scheduling Conflicts 75 99 Regulations 3 35 Incorrect Algorithm Design 55 72 Inconsistent Motor Construction 22 57 Environment Prevents Recovery 25 75 Loss of Source Code 50 52 Poor Quality Presentation 10 33 Poor Weather 18.5 29 Scheduling Conflict With The Schools 12 66 Burn Ban in Effect 25 42 Power Budget Miscalculation 1 92 Loss of Rocket Lab 5 Loss of Low-Altitude Test Launch 93 Facility in Glen Rose, TX 5 Loss of High-Altitude Test Launch 94 Facility in Cross Plains, TX 5 38 Loss of Science Building 0 39 Engineering Building 0 40 Loss of HCI facilities 12 Table 12 Project Risks Tarleton Aeronautical Team 43 2012 - 2013 USLI Preliminary Design Review

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

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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

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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.

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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

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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.

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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

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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

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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.

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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.

Unit Unit Total Item Distributor Product Number Dimensions Cost Cost

Main Featherweight 1.8in. X 0.8in. X Raven3 $155.00 2 $310.00 Altimeter Altimeters 0.55in. X 0.34oz

3.75in. long X Loki Research Ozark ARTS 1.4in. wide X $190.00 2 $380.00 2.75 oz 2.75in. long X Backup PerfectFlite Stratologger 0.9 in. wide X $79.95 2 $159.90 Altimeter 0.45 oz 0.9in. X 0.65in. X Adept Rocketry ALTS1-50K $89.00 2 $178.00 4.25in. X 4.25oz Table 18 Deployment Altimeter Trade and Selection

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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

Figure 19 Daveyfire Electric Match

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

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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.

Unit Pre- Unit Distributor Item Dimensions Assembled Cost Coast Daveyfire N28BR 1 ft long X 1 Yes $2.95 Rocketry QuickBurst E-Match QuickBurst 1 ft long X 20 No $32.00 Kit 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,

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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

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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

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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.

Wind Mount Unit Speed Water Unit Component Distributor Item for Dimensions and Proof Cost Tripod Direction

4500 Pocket Ambient Kestrel 5in. X 1.8in. Weather Yes Yes Yes $299.00 Weather 4500 X 1.1in. Tracker

Speed 5.5in. X WindMate Weather Tech 1.75in. X Yes Yes Yes $229.95 Anemometer Shack WM- 0.75in. 350

Speed 5.5in. X WindMate Weather Tech 1.75in. X Yes Yes Yes $154.95 Anemometer Shack WM- 0.75in. 300 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.

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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:

( )

( )

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∑

.

( )

( ) ( )

√

√

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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√

√ ( )

The calculated descent rate of the vehicle after main parachute deployment at 500 feet is 12.7246 feet per second. With this figure it possible to calculate the total drift of the vehicle after main parachute deployment until landing. The following math calculates the drift of the vehicle from main parachute deployment until landing assuming a 15 mile per hour (22 feet per second) horizontal wind velocity.

With a 15 mile per hour wind, the vehicle will drift approximately 864 feet after main deployment at an altitude of 500 feet, until landing. Since the maximum allowable drift is 2500 feet, the drift between apogee and the main parachute deployment can only be 1636 feet, 2500 minus 864 feet. The following calculations give the minimum descent rate for the vehicle while the drogue parachute is deployed, taking into account a drift of less than 1636 feet.

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Thus the minimum descent rate from apogee to main deployment in order to keep total drift below 2500 feet is 65.8924 feet per second. With this figure it is possible to calculate a maximum drogue parachute size that causes the descent rate to be less than 65.8924 feet per second. The following calculates the maximum diameter of the drogue parachute:

( )

( )

∑

.

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( )

( ) ( )

√

√

Based on the above calculations, commercial availability, and an overestimate of 5400 feet for apogee, the selected drogue parachute diameter is 24 inches.

Parachute Diameter Drogue 2 ft Main 10 ft Table 22 Parachute Diameters

The parachutes are 1.1 ounce silicon-coated nylon to ensure durability and strength. Basing information on the resilient construction of the parachutes, flight simulations from OpenRocket (see table 26), and commercially available size limits, the diameter of the drogue is 2 feet, while that of the main 10 feet.

Recovery Component Itemization

Parachute Selection Of the commercially available parachutes in consideration, those which made it

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into the trade and selection process have the least number of shroud lines to limit the probability of entanglement upon ejection, as well as reinforced seams with nylon webbing to reduce the probability of shroud line disconnection as a result of ejection. The selections of the round nylon parachutes for each the main and drogue parachutes is in Table 23 and Table 24.

Shroud Unit Manufacturer Item Diameter Lines Cost Top Flight TFR Par-120 16 10ft $89.95 Recovery STANDARD LOW-POROSITY 1.1 $110.0 The Rocketman 4 10ft RIPSTOP PARACHUTE 0 $240.0 Fruity Chutes Custom Parachute - 50lb @ 20fps 16 10ft 0 Table 23 - Main Parachute Trade and Selection

Shroud Diamete Unit Manufacturer Item Lines r Cost Top Flight PAR-STD-24 6 2ft $9.95 The STANDARD LOW-POROSITY 1.1 4 2ft $25.00 Rocketman RIPSTOP Apogee Nylon Parachute #29218 6 2ft $11.44 Rockets Table 24 - Drogue Parachute Trade and Selection

Parachute Protection Each parachute is quick-linked to the bridle of a 2,000 degrees Fahrenheit-plus- heat-resistant Nomex deployment bag. Cellulose (“dog barf”) insulation and sheet insulation wadding were also viable candidates for their wide availability and low cost. Through testing, it is determinable whether dog barf is usable in conjunction with the deployment bag system.

However, hand-packed insulation such as dog barf or sheet wadding could easily shift during ascent, potentially exposing some portion of the parachute, shroud lines, or shock cord. This could result in a fire hazard and entanglement. The deployment bag keeps the parachutes, shroud lines, and shock cord away from the heat of the firing ejection charge reliably throughout ascent and allows for ejection in an orderly fashion. Table 25 demonstrates this selection.

Unit Component Distributor Item Unit Dimensions Cost

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Flameproof Main Parachute Rocketman 3.25in. X 3.25in. X Deployment Bag Enterprises DB8 6.5in. (custom) $40.00 Recovery Wadding for Main Compartment HobbyLinc EST302274 75 sheets $3.69 Dog Barf for Main Rockets "R" "Dog Barf" Recovery Compartment Us Wadding - 16oz 16oz $7.00 Flame-Proof Drogue Rocketman 3.25in. X 3.25in. X Parachute Deployment Bag Enterprises DB2 18.5in. (custom) $25.00 Recovery Wadding for Drogue Compartment HobbyLinc EST302274 75 sheets $3.69 Dog Barf for Drogue Rockets "R" "Dog Barf" Recovery Compartment Us Wadding - 16oz 16oz $7.00 Table 25 - Parachute Protection Materials

It is important to know that these parachutes have a probability of entanglement, as the main parachute has 16 shroud lines; the drogue has six. To counteract this issue, standard high-power model rocketry 900 pound-working-strength swivels are on each end of each shock harness, connecting each shock harness to both the parachute and the eyebolt. Each parachute is neatly folded into its deployment bag with all shroud lines and shock cord. This keeps each parachute inside long enough to maintain separation from the fins and to ensure that ejection takes place in an orderly fashion.

Shock Chord

In order to enable deceleration of the two tethered sections on descent, 0.5-inch tubular Kevlar shock harnesses are necessary. These implementations are for strength and flame-proof construction. Additionally, these shock harnesses include pre-sewn Nomex loops for a safe and secure connection to the parachutes by delta-shaped quick links.

The main harness is to be approximately two body lengths plus fifteen percent to keep the main parachute away from the body. The drogue harness is to be approximately three body lengths plus fifteen percent to pull the drogue parachute up and away from the vehicle, while minimizing the risk of denting or zippering. The body length is approximately nine feet, including the nose cone. Thus, based on commercial availability, the length of the main shock harness is to be 20 feet, while that for the drogue is to be 25 feet. Table 26 illustrates this selection.

Unit Unit Component Distributor Item Dimensions Cost Main Shock Cord Giant Leap Rocketry Tubular Kevlar 0.5in. X 25ft $37.99

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Rocketman Bulk Tubular Nylon 1in. X 20ft $30.00 Enterprises Webbing Drogue Shock Giant Leap Rocketry Tubular Kevlar 0.5in. X 20ft $31.49 Cord Rocketman Bulk Tubular Nylon 1in. X 15ft $25.00 Enterprises Webbing Table 26 - Shock Chord Trade and Selection

Stainless steel delta-shaped quick links, each with a working load of 1,000 pounds, are a in the design. These secure each parachute to its shock harness and each shock harness to a U-bolt at the bulkhead of each altimeter compartment to ensure secure tethering throughout the duration of the flight.

Shear Pins In order to ensure that separation in the rocket body only occurs upon ejection, the design makes use of removable threaded nylon shear pins. These are inserted through holes on either side of the couplings between the nose cone shoulder and main parachute compartment, as well as between the payload and drogue parachute compartments. When the ejection charge fires, the force of the coupler sliding past will snap the shear pins. However, other stresses under 25 pound force such as those caused by shifting mass, drag, or ejection from another compartment should not be strong enough to cause separation.

The itemized components of the recovery section are in the following Table 27. This includes the total cost and mass (where available) of each item. The total cost reflects one fully-assembled rocket.

Table 26 – Launch-Day Recovery System Budget Proposed Unit Unit Qty. Distributor Item Number Total Cost Selection Dimensions Cost

1.8in. X 0.8in. Main Featherweight Raven3 X 0.55in. X $155.00 2 $310.00 Altimeters Altimeters 0.34oz

2.75in. X Backup PerfectFlite StratoLogger 0.9in. 0.55in., $79.95 2 $159.90 Altimeters 0.45oz

Electric Daveyfire Coast Rocketry 1ft long $2.95 4 $11.80 Matches N28BR

FFFFg Black Goex 4F Black Goex 1lb $15.75 1 $15.75 Powder Powder

Black Powder Aerocon Ejection BPSmall 15 X 0.067oz $3.00 1 $3.00 Systems Charge

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Holders

Commonwealth Swivels SWLDK80 1.5in. $1.99 2 $3.98 Rocketry

Main Shock Giant Leap Tubular Kevlar 0.5in. X 25ft $37.99 1 $37.99 Cord Rocketry

Drogue Shock Giant Leap Tubular Kevlar 0.5in. X 20ft $31.49 1 $31.49 Cord Rocketry

Main Top Flight TFR Par-120 10ft Diameter $89.95 1 $89.95 Parachute Recovery

Flameproof 3.25in. X Main Rocketman 3.25in. X Parachute DB8 $40.00 1 $40.00 Enterprises 6.5in. Deployment (custom) Bag

Drogue Top Flight PAR-STD-24 2ft Diameter $9.95 1 $9.95 Parachute

Flame-Proof 3.25in. X Drogue Rocketman 3.25in. X Parachute DB2 $25.00 1 $25.00 Enterprises 18.5in. Deployment (custom) Bag

U-Bolt 0.212in. X Sunward Assembly - U-Bolts 1.5in. X 2.125 $4.29 2 $8.58 Aerospace 0.25in. X 1.77oz (compact)

0.25in. 2.375in. X Commonwealth Stainless Steel Quick Links 0.375in. X $2.99 4 $11.96 Rocketry Delta Quick 1.25in. Link

2-56 Nylon 0.08in. X Shear Pins Missile Works Shear-Pin (10 0.5in. X $1.00 1 $1.00 pack) 0.002oz X 10

Featherweight Arming Featherweight 0.55in. X Magnetic $25.00 2 $50.00 Switches Altimeters 0.75in. Switch

Hand-held SpeedTech 5.5in. X 1.7in. Rotary Fan Weather Shack $154.95 1 $154.95 WM-300 X .75in. Anemometer

5.5in. X BeeLine GPS- GPS System Big Red Bee 1.75in. X $289.00 2 $578.00 Package Deal 0.75in.

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Radio Pure Copper Frequency 12in. X 42.5in. LessEMF Polyester $10.95 2 $21.90 Shielding X 0.003 in. Taffeta Fabric Material

Total $1,565.20 Table 27 Launch Day Recovery System Budget

Mission Performance Predictions

Mission Performance Criteria

The rocket vehicle delivers the SMD payload to 5,280 feet above ground level. It carries a barometric altimeter for official scoring, and must remain subsonic during the entire flight, from launch until landing. The vehicle must be recoverable and reusable on the day of the official launch.

Simulations Flight Simulations

Performances of flight simulations in OpenRocket help gather data concerning expectations for the vehicle. All relevant parameters are input into this simulation to acquire vehicle performance data including component weights. The mass of the rocket is slightly over-estimated to account for weight gain as the design progresses. The simulations account for launch conditions using a yearly average for the date and location of the official launch site. Launch conditions include 9±.9 mph average wind speed, where ±.9 mph was found using 10% turbulence intensity. A 75° F temperature and 1241 mbar pressure is in place for the simulation as well. All weather data accumulation is from the National Oceanic and Atmospheric Administration (NOAA). Latitude and longitude locations of 34.9° N and 86.5°W are applicable for the project.

Altitude

With these launch conditions, the vehicle achieves a height of 5350±5 feet. The following graphs illustrate the altitude throughout the simulated flight. After 2.11 seconds of motor burn, the estimate for altitude of the vehicle is 760 feet. The

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vehicle continues to climb to apogee at 5,245 feet where the initial separation event and drogue parachute deployment occur. This precedes the main parachute deployment at 500 feet.

Figure 20 Altitude Simulation

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Velocity

Under guidance from the team mentor, Pat Gordzelik, an exit velocity over 60 feet per second is the goal to ensure adequate stability when leaving the launch rail. Additionally, the intent is for the vehicle to remain subsonic to meet the requirements set forth in the Statement of Work (SOW). The following graph shows the entire simulation of the velocity from launch until landing. After 0.24 seconds of motor burn, the velocity of the vehicle when departing from the launch rail should be 76 feet per second; this exceeds the recommended 60 feet per second. After a total of 2.11 seconds when the motor propellant depletes, a maximum vertical velocity of 833 feet per second is the result; this demonstrates that the vehicle remains subsonic throughout the flight. Apogee should occur after 17.68 seconds into the flight, with a theoretical altitude of 5,345 feet. Drogue parachute deployment occurs at apogee and the vehicle begins to descend at a velocity of approximately 75 feet per second until the main parachute deployment occurs at 500 feet. At this time, the velocity decreases to 17.7 feet per second until landing.

Figure 21 Velocity Simulation

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Acceleration

Throughout the flight, multiple stages of acceleration occur. Two graphs illustrate the acceleration of the vehicle throughout the flight. The first stage represents the acceleration from ignition to burnout of the motor. The maximum acceleration of the motor is approximately 400 feet per second per second. This occurs 0.789 seconds into the flight. At approximately 2.11 seconds into the flight, just prior to motor burnout, the acceleration decreases.

Figure 22 Acceleration Before Burn Out Simulation

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The second graph represents the simulated acceleration from motor burn out to landing. At apogee, the acceleration is -31.28 feet per squared second. At an altitude of 500 feet, the main parachute deploys, and the vehicle momentarily accelerates to 553 feet per squared second for less than .1 second. The acceleration then falls into an oscillation, ranging from -.15 to .15 feet per squared second for the remainder of the descent to ground level.

Figure 23 Acceleration After Burn Out

Figure 24 is the simulated thrust curve for the Cesaroni L1720-WT motor.

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Figure 24 L1720 Thrust Curve

Stability Center of Pressure/Gravity

The stability margin should be at least two calibers of body diameter. The center of gravity is 65.667 inches from the nose, while the center of pressure is at 79.738 inches from the nose. This provides 14.071 inches of stability margin or 2.55 calibers of body width.

Figure 25 Stability: Center of Pressure/Gravity

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Kinetic Energy

The weight of the rocket upon landing is 29.62 pounds. The weight of the nose cone section is approximately 2.252 pounds, while that of the upper body airframe and payload is approximately 16.766 pounds, and the booster section is approximately 10.602 pounds, excluding propellant.

The following calculates the kinetic energy, first for the entire rocket, then for each individual section. First it is necessary to calculate the descent velocity upon landing.

√

√ ( )

Using 12.7246 feet per second for the landing velocity it is possible to calculate the kinetic energy for the entire vehicle and each individual section, with relation to their mass. The following calculates the kinetic energy of the entire vehicle.

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( )

( )

Therefore the kinetic energy for the entire vehicle as a whole is 74.471 foot pound force. This leads to the assumption that if the kinetic energy of the entire vehicle is 74.471 foot pound force, then the kinetic energy for each individual section must be less than 74.471 foot pound force. It is possible to prove this assumption by calculating the kinetic energy of each individual section. The following calculates the kinetic energy of the nose cone, which has a weight of 2.252 pounds.

( )

( )

The following math calculates the kinetic energy of the connected upper body air frame and payload. This section has a weight of 16.766 pounds.

( )

( )

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The following calculates the kinetic energy of the booster section. This section has a weight of 10.602 pounds.

( )

( )

Table 28 summarizes the kinetic energy for each individual section and the vehicle as a whole.

Section Weight Kinetic Energy Nose Cone 2.252 lbs 5.662 ft∙lbf Upperbody Airframe and Payload 16.766 lbs 42.1533 ft∙lbf Booster 10.602 lbs 26.6557 ft∙lbf Total 29.62 lbs 74.471 ft∙lbf Table 28 Kinetic Energy Summarization Drift Calculations Wind causes two effects during flight, drift after parachute deployment and weather cocking during vertical ascent. The parachute selection calculations do not take in to account weather cocking but rather calculate the drift after parachute deployment. The following formula is the basis for the drift calculations. The formula takes in to account the two different descent rates, one for when the drogue parachute is deployed and one for when the main parachute is deployed. The apogee altitude is approximated to 5280 feet.

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The following calculates the theoretical landing radius of the vehicle for zero miles per hour. These results do not take into account weather cocking.

For five miles per hour (7.33 feet per second):

For ten miles per hour (14.67 feet per second):

For fifteen miles per hour (22 feet per second):

( )

For twenty miles per hour wind (29.33 feet per second):

The previous results did not take into account weather cocking. Weather cocking occurs when a horizontal wind velocity rotates the vehicle to where it has a new flight direction into the wind. Figure 26, found on the NASA website, demonstrates the effects of weather cocking.

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Figure 26 Weather Cocking

Weather cocking causes the vehicle to fly into the wind. Then after apogee, horizontal wind velocity causes the vehicle to drift with the wind. This effect greatly decreases the landing radius when compared to theoretical calculations. Open Rocket simulations calculate weather cocking in flight trajectory.

Figure 27 is a plot of the vehicle’s distance from the launch pad in relation to altitude. The simulation assumes a 15 mile per hour wind.

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Figure 27 Weather Cocking Simulation

Table 29 summarizes the simulated landing radius of the vehicle for five different wind speeds. These calculations take in to account both weather cocking and parachute drift.

Wind Speed Landing Radius 0 mph 7 ft 5 mph (7.33 ft/s) 275 ft 10 mph (14.67 ft/s) 550 ft 15 mph (22 ft/s) 900 ft 20 mph (29.33 ft/s) 1300 ft Table 29 Landing Radius

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Interfaces and Integration

Payload Integration Plan

The payload is designed to integrate into the payload housing structure of the launch vehicle in a simple and easy fashion. The payload is constructed on a payload framework which consists of two bulkheads connected to each other by two aluminum rails. Because the payload housing structure is part of the vehicle, the payload framework uses dimensions to ensure compatibility between the payload and the vehicle.

Figure 28 Payload Pre-Integration

The payload is built onto the payload framework, and preparing the payload for flight is easily done prior to payload framework integration with the vehicle. At this time the framework is inserted from the lower portion of the payload housing structure. Once the payload framework is in place, it is secured with the use of two screws; from the exterior of the payload housing structure and into the bulkhead. At this point the physical integration of payload is complete.

Internal Vehicle Interfaces

Structure

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Structure(heading lv 4) (upper body structure=upper body airframe; lower body structure= booster section)

The flight vehicle consists of four main sections; the nose cone, the upper body airframe, the payload housing structure, and the booster section. These sections interface with couplers, rivets, screws, and shear pins. Two identical 11 inch long couplers are modified to have a larger outer diameter on one side. This modification allows compatibility of the inner diameter of the acrylic payload housing structure. Each coupler interfaces the payload housing structure with the fiberglass upper body airframe and booster section respectively. Figure 29 shows the coupler specifications.

Figure 29 Coupler Specifications

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The nose cone shoulder is used to couple the nose cone with the upper body airframe. This coupling is secured with shear pins, as the nose cone must separate from the upper body airframe to deploy the main parachute. A Kevlar shock cord is used to tether the nose cone and the upper body airframe. The Kevlar tether is attached between two eyebolts.

The upper body airframe and the payload housing structure are coupled with a fiberglass coupler and four aluminum rivets; two rivets through the upper body airframe, and two additional rivets through the payload housing structure. The rivets attaching the payload housing structure with the coupler will remain unaltered for the life of the vehicle; whereas, the rivets connecting the upper body airframe to the coupler may be drilled out and reinstalled to gain access to the components in the lower portion of the upper body airframe, namely the GPS bay and main parachute altimeter bay.

The payload housing structure, from the bottom end, couples with the booster section using a fiberglass coupler, four screws, and two shear pins. The screws will attach the bottom portion of the payload housing structure to the fiberglass coupler and the bottom bulkhead of the payload framework. The booster section and the coupler attach as a friction fit, but also utilize two shear pins. Additionally there is an eyebolt anchored at the bottom of the payload framework bulkhead, which secure to the payload housing structure. A Kevlar shock chord tethers the payload housing structure eyebolt to another eyebolt in the booster section.

For the booster section, fins mount through slits in the vehicle body and epoxy in place. The motor mount tube attaches with centering rings to the inner surface of the booster section and epoxy in place.

Launch Vehicle to Ground Interfaces

The recovery system altimeter arming switches are a mechanical interface with the ground. Physically removable magnetic pins activate the arming switches upon removal. The pins are not removed until the vehicle is on the launch pad and ready for flight. Upon removal, the altimeters audibly respond with arming confirmation. The drogue deployment altimeter arming switch is located two feet above the base of the vehicle and the main deployment altimeter is located five feet eight inches above the base of the launch vehicle, which does not exceed six feet (Requirement 2.9). The BeeLine GPS and the XBee XSC S3B radio, payload section, are wireless transmission interfaces between the ground and the vehicle. Neither wireless device receive telemetry, but both transmit telemetry to the

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ground station. The BeeLine GPS transmits readings of latitude and longitude for recovery purposes and the XBee XSC s3B transmits atmospheric data. Table 30 summarizes the interfaces between the vehicle and the ground.

Ground – Vehicle Interface Device

Mechanical Arming Switches

Wireless Transmission BeeLine GPS, Xbee XSC S3B Table 30 Ground- Vehicle Interface

Vehicle to Ground Launch System Interfaces

Rail buttons are attached to the vehicle to provide an interface between the vehicle and the launch rail. The selected rail buttons function with a 1010 launch rail. The bottom rail button is located two inches from the base of the vehicle and the upper rail button is located fifteen inches from the base of the vehicle. The motor ignition system interfaces the vehicle and the ground launch system. The motor ignites by a standard twelve volt firing system. The igniter is provided with the Cesaroni L1720-WT-P motor.

Launch Operation Procedures

Appendix B contains checklists for the final assembly and launch procedures. The checklists provide a detailed description of the steps to ensure a successful flight. The checklists will be strictly followed and updated as needed.

Safety and Environment (Vehicle)

The Safety Officer

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The team safety officer, Blake, is level one certified with NAR. The responsibility of the safety officer is to design and implement safety plans that ensure all accidents are evaded. All hazards to people, the project, and the mission are determined so that mitigations can be enacted.

The systematic identification of risks, failure modes, and personnel hazards allows for the team to discover where single points of failure could occur throughout the course of the project. The identification of single point failures allows for proactive design changes to be made to counter these failures.

Failure Modes

A failure mode is the way in which a system could fail, causing an undesirable effect on some aspect of the project. The safety plan ensures development and implementation of mitigations for each failure mode. Rocket Design Failure Modes Table 31 provides a summary of potential failure modes that could occur during the design of the vehicle.

Failure Effect Proposed Mitigation Completed Mitigation Unpredictable Vehicle Unstable Simulations Completed Flight Path

Flight Testing Proposed (12/1/2012)

Acrylics Does Not Vehicle Not Tensile Strength and Withstand Forces Proposed (12/1/2012) Reusable Flight Testing Throughout Flight Fiberglass Does not Withstand Vehicle Not Tensile Strength and Proposed (12/1/2012) Forces Reusable Flight Testing Throughout Flight Connection Between Acrylic Unpredictable Payload Housing Flight Path, and Upper Research/ Design Completed Damage to Fiberglass Body Vehicle Body Tube Becomes Detached

Testing Proposed (12/1/2012)

Expected Fins Cause To Apogee Height Simulations Completed Much Drag Not Obtained

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Flight Testing Proposed (12/1/2012)

Thrust To Weight Unpredictable Simulations/ Ratio is Less Than Completed Flight Path Calculations 5:1 Research/ Couplers Too Early or No Simulations/ Completed Long or Too Short Separation Calculations

Flight Testing Proposed (12/1/2012)

Table 31 Potential Failure Modes for the Design of the Vehicle

Payload Integration Failure Modes Table 32 provides a summary of potential failure modes that could occur during payload integration.

Failure Effect Proposed Mitigation Completed Mitigation Inadequately Ensure the bulkhead is replaced Screw hole Proposed (1/5/2012) stripped out secured payload when required Excessive Be sure to replace when Rubber O-ring Proposed (1/5/2012) tear vibrations required Circuitry Electronic Be careful while inserting becomes hung Proposed (1/5/2012) malfunction payload up Table 32 Potential Failure Modes during Payload Integration Launch Operations Failure Modes Table 33 provides a summary of potential failure modes that could occur during launch operations.

Potential Failure Potential Effects of Proposed Completed Subsystem Mode Failure Mitigations Mitigations Ejection charges damage air- Critical systems Calculations Completed frame/rocket become damaged Structure components Proposed Proper testing (1/5/2012) Motor mount fails to Structural testing Damage to internal Proposed properly retain of the motor systems (10/27/2012 motor mount Ensure rail Unpredictable flight buttons are Proposed Rail button failure path properly installed (1/5/2012) and orientated

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Insufficient Potential system Test mounting Proposed component malfunction integrity, (1/5/2012) mounting Loss of vehicle Airframe stress Structural testing Proposed functionality, potential failure of airframe (1/5/2012) loss of vehicle Ensure that the Aerodynamic Proposed Fin Detachment fins are properly instability of vehicle (12/1/2012) epoxied Potential Failure Potential Effects of Proposed Completed

Mode Failure Mitigation Mitigations Parachute shroud Uncontrollable Research Completed line fails descent Verify parachute Proposed rating (10/27/2012) Parachute Electronic matches Redundant deployment does not Completed do not fire altimeter system occur Uncontrollable Verify eye bolt Proposed Eye bolt failure descent integrity (1/5/2012) Untethered vehicle Properly fastened Proposed Shock cord failure components, violation shock cord (10/727/2012) of requirements Proposed Recovery Verify rating (10/27/2012) Premature black Premature parachute Proposed Testing power ignition ejection (10/27/2012) Recovery Completed altimeter shielding

Failure of recovery Use nomex cloth Recovery system Proposed system, damage or and fire retardant ignites (10/27/2012) loss to vehicle insulation Main or drogue Secure swivels parachute comes Proposed Uncontrolled descent along with quick untied from the (1/5/2012) links. swivel Ensure parachute Main or drogue Uncontrolled descent shroud lines are Proposed parachute shrouds rate attached to a (1/5/2012) become entangled swivel Un-functional Adequate Proposed Failed Separation recovery system, separation testing (10/27/2012) ballistic descent failure of deployment PerfectFlites power Use new batteries Proposed of parachutes, supply diminishes before launch (1/5/2012) Mission failure Featherweight failure of deployment Use new batteries Proposed Power supply of parachutes, before launch (1/5/2012) diminishes Mission failure PerfectFlite wired failure of deployment Use protected connections of parachutes, electrical Completed become damaged Mission failure components

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from handling

Featherweight failure of deployment Use protected wired connections of parachutes, electrical Completed become damaged Mission failure components from handling We lose all points Scoring Altimeter associated with the Redundant Completed failure altitude portion of the Systems project Potential Failure Potential Effects of Proposed Completed Mode Failure Mitigation Mitigations Igniter does not initiate the oxidation The rocket does not Inspect igniter for Proposed process for the launch concatenation (12/1/2012) propellant Always bring Proposed additional igniters (12/1/2012) for such an event Use proper Propellant’s igniter, sue oxidation process The rocket does not Proposed appropriate does not launch (12/1/2012) conditions when commence Propulsion storing propellant A pressure build-up Proposed occurs inside the Explosion Inspect the motor (12/1/2012) motor Table 33 Potential Failure Modes during Launch

Hazard Analysis

Table 34 provides an overview of potential hazards to personnel through the course of the project. Personnel hazards refer to potential harm incurred by any individual. The development and implementation of the safety plan and protocols ensure that these hazards are appropriately mitigated.

Risk Sources Likelihood Consequences Mitigation Action

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Follow safety protocols, proper tool Knives, and Discontinue all Laceration routers, Serious injury equipment operations, Medium saws, file, or death use, apply first aid, Dremel tool personal contact EMS safety attire, refer to operators manual Follow safety protocols, Chemicals proper tool (FFFFg, and Discontinue all fiberglass Burns Minor to serious equipment operations, resin), Medium injury use, apply first aid, welders, personal contact EMS soldering safety attire, Iron refer to operators manual Follow safety protocols, Respiratory Chemicals proper tool Discontinue all Damage (epoxy, and Brain damage operations, solder), Low equipment or death apply first aid, fumes, use, contact EMS fiberglass personal safety attire, consult MSDS Use of goggles, force Welders, shields, Discontinue all Vision Damage fiberglass, Partial to consult operations, grinders, Low complete MSDS, first apply first aid, projectile blindness aid kit contact EMS, debris available, use eyewash refer to operators manual Discontinue all Use of Loss of operations, gloves, Epoxy, respiration, apply first aid, consult Allergic chemicals, Low inflammation contact EMS, MSDS, first Reaction fiberglass (Internal & administer aid kit External) antihistamines, available safety shower

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Ear muffs, consult FFFFg, MSDS, first Discontinue all Partial to Hearing Grinders, aid kit operations, Low complete Damage Ignition, available, apply first aid, deafness Routers refer to contact EMS operators manual Make sure Discontinue all proper operations, Projectiles, safety Permanent apply first aid, Dismemberment Saws, Low measures injury or death and contact Launches are taken, EMS, operators tourniquet manual Table 34 Potential Hazards to Personnel

The material safety data sheet (MSDS) that the manufacturer provides contains pertinent information about the material in consideration. It is comprised of sixteen categories: identification, hazard(s) identification, composition/information on ingredients, first-aid measures, fire-fighting measures, accidental release measures, handling and storage, exposure controls/protection, physical and chemical properties, stability and reactivity, toxicological information, ecological information, disposal information, transport information, and regulatory information. MSDSs are referred to when a hazard occurs in order to enact the most effective mitigation. All team members shall be knowledgeable of the MSDS associated with each hazardous material. Operator manuals for each tool will be consistently referenced prior to each tool’s usage. This ensures each tool is used as intended. According to the safety plan, operator manuals for each tool used during the project will accompany the MSDSs in the safety binder. These documents will be made available by the safety officer at any location in which construction, testing, or launching of the vehicle could occur.

It is important for all team members to be thoroughly briefed on the project risks, FAA laws and regulations regarding the use of airspace, and the NAR high-power safety code.

The team is aware that the FAA must be notified of planned launch activities. For educational outreach events, notification to the closest airport within 5 miles of the launch site is required 72 hours prior to launch. For subscale launches, flight waivers are required to be obtained at least 45 days prior to the proposed activity. For full scale launches, access to launch sites provided by West Texas Rocketry Tripoli number 121 will be utilized.

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The National Association of Rocketry and Tripoli are recognized as the primary rocketry associations of the United States. As such, their standards establish precedence throughout high powered model rocketry. Along with these standards, the team is cognizant of all federal, state, and local laws regarding unmanned rocket launches and motor handling, including the following regulations: CFR 101, Subchapter F, Subpart C: Amateur Rockets (Located in Appendix D) CFR Part 55: Commerce in Explosives (Located Appendix D) Handling and Use of Low-explosives Ammonium Perchlorate Rocket Motors (APCP) (Located in Appendix I.12) NAR Model Rocket Safety Code (Located in Appendix F) Hazardous Waste Management (Located in Appendix E) Fire Safety (Located in Appendix G) Lab Safety (Located in Appendix H)

Table 35 provides a summary of legal risks that could occur during the course of the project.

Risk Likelihood Severity Consequence Mitigation FAA Violations Legal Adhering to Low High Repercussions Regulations NAR/TRA Legal Adhering to Low High Violations Repercussions Regulations Damage of Legal Property Low High Insurance Repercussions

OSHA Legal Adhering to Low High Violations Repercussions Regulations Personal Redundant Injury Legal Calculations Low High Repercussions and Safety Preparedness Table 35 Summary of Legal Risks Environment Environmental effects of the project

In the event of an unrecoverable or damaged rocket, certain materials could be left exposed to the environment. The biodegradability of each material used affects the impact on the surrounding ecosystem. Much of the information on the hazards posed to the environment and ecology is available in the individual MSDSs. The

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effects of materials used in the construction and launch of the rocket are summarized in the following table 36.

Mode of Material Prevalence Impact on Environment Biodegradability Ammonium Iodization of local water Motor propellant Highly water soluble Perchlorate table Black Powder Ejection charges Remains solid No known impact Fiberglass Decomposition begins Leaching to local water connections, sealed within fifteen months table Epoxy 2M DP420 couplings External fiberglass Leaching to local water Oil-Based Spray structural Soluble table Paint: components Leaching to local water Payload bay Soluble Clear Acrylic: table No known Structural environmental impact; Remains solid components may pose ecological Fiberglass hazard No known environmental impact; Deployment bag Remains solid may pose ecological Nomex hazard Motor tube, rivets, Long-term degradation Highly water soluble Aluminum battery casing, products Long-term degradation Recovery wadding; Remains solid Cellulose products Attachment Leaching to local water hardware, ballast Remains solid table Steel system; Avionics bay lining, Highly reactive in air or Long-term degradation Copper: e-match lead wires moisture products Long-term degradation Batteries Highly water soluble products; may pose Sulfuric Acid ecological hazard Kevlar Shock harnesses Remains solid No known impact Parachutes, Reactive in air or moisture Irritating vapors form Silicon batteries Parachutes, shear Remains solid No known impact Rip-Stop Nylon: pins Table 36 Effects of Materials used in Construction and Launch

Environmental effect on the project

While some aspects of the project may adversely affect the surrounding environment, the environment can also have an impact upon the project. As primary test launches will take place in Texas, during winter and spring, inclement weather will likely fall on test dates. In response to unforeseen issues in the

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weather, alternate test dates have been scheduled and can be viewed in the testing timeline in Figure 62. Environmental factors such as surrounding flora and fauna or sedimentary projections could cause the launch vehicle to become unrecoverable. These risks are outlined in Table 37.

Risk Likelihood Severity Consequence Mitigation Poor weather Multiple test High High Delay in testing dates Environment Survey launch prevents Possible loss of Medium Medium site, recovery recovery Vehicle tools Burn ban in effect Multiple test Low High Delay in testing locations

Table 37 Environmental Factors

IV) Payload Criteria Selection, Design, and Verification of Payload Experiment

System Level Review At a system level, the payload design fulfills the requirements of the SMD payload. This payload option was selected because of the high difficulty level. The Tarleton Aeronautical Team is honored to be able to apply personal skills and knowledge to a real world scientific problem. The payload consists of redundant sensors and power supplies. The payload electronics sub-team will implement the incremental development process. At the completion of each increment, a fully functioning product is developed. The deadline for the first increment is November 30th. The deadline for the second increment is December 30th. During the first increment the electronics are mounted on vertical aluminum rails. Connections between electronics are assisted by perforated circuit boards. A model of the first increment product is shown in Figure 30. After prototyping is successful, the next increment of the development process includes fully integrating printed circuit boards (PCBs). This reduces weight and size requirements of the payload.

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Figure 30 Payload Design Configuration

The payload gathers measurements of pressure, temperature, relative humidity, solar irradiance, and ultraviolent radiation (Requirement 3.1.3.1). Several readily available sensors, along with a microcontroller equipped with the appropriate number of digital and analog inputs are selected. Redundant barometric pressure sensors, redundant temperature sensors, redundant humidity sensors, redundant UV radiation sensors, and redundant pyranometers are included in the design. Measuring UV radiation and solar irradiance continuously throughout the flight is achieved by placing a UV sensors and pyranometers facing every ninety degrees. This design is shown in Figure 30. For consistency in data and sophistication in design purposes the plan is to research an autonomous solar tracking system for the UV radiation sensor and the pyranometer. The chosen microcontroller contains the necessary data busses to communicate with all sensors.

The chosen microcontroller utilizes a microprocessor with a processing clock speed capable of handling all the necessary sensors at the required frequencies (Requirements 3.1.3.2 and 3.1.3.3). The clock speed of the microcontroller’s microprocessor is 16 megahertz. Surface data collection terminates ten minutes after landing (Requirement 3.1.3.4). The microcontroller is able to record the time in milliseconds since the start of the most current software program. The flight software will use sensor data to detect landing, note the current time in milliseconds since the program started, then compare the current time to the time at landing for every loop cycle; when the current time is 600,000 milliseconds (ten minutes) larger than the time at landing, surface data collection will terminate.

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The payload includes the ability to take two pictures during descent and three pictures after landing, while keeping the orientation of the pictures with the sky towards the top and the ground towards the bottom of the frame (Requirements 3.1.3.5 and 3.1.3.6). To ensure this functionality, our payload contains a camera with a dedicated microcontroller, data storage device, and power supply. Several options for camera orientation are considered. One option is to mount the camera on a pendulum connected to a ball bearing. This method utilizes gravity and was very efficient due to no power requirements. However, considering the orientation of the payload, oscillations and rotational motion on descent is a major setback to this option. The pendulum also has to be kept extremely stable and limited in its range of motion; otherwise the camera could swing into the wall of the payload housing.

Another alternative is to place a camera on each side of the payload, opposed 180 degrees. Throughout the flight a multitude of pictures would be taken, and during post flight analysis the orientation of the pictures would be adjusted. This design follows the Keep It Simple (KIS) principle; however, the team wants a more reliable and sophisticated solution. The current design has the camera mounted on servo motors. The correction angle of each servo is determined by an accelerometer measuring the tilt of the payload. By measuring the change in orientation of the payload, the servo motors can correct the camera’s position such that pictures are captured in proper orientation. Although this design requires a large power supply and an additional sensor, the overall sophistication and reliability of the design are the determining factors for selection of this option. Figure 31 demonstrates the physical layout of the self-leveling camera.

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Figure 31 Self-leveling Camera Configuration

Payload data is stored onboard and transmitted wirelessly to the ground station (Requirement 3.1.3.7). A 16 gigabyte high speed data storage medium is selected. Each measurement taken by the sensors along with any other usable data is saved to this data storage medium. A 900 megahertz wireless transmitter incorporates into the payload electronics. The microcontroller assembles a telemetry string of sensor data. The sensor data is delimited by space characters. The ground station includes a wireless receiver connected to a handheld high gain Yagi antenna. A MATLAB graphical user interface program parses the telemetry string, displays the readings, and saves each telemetry string to a file.

The payload carries a GPS tracking unit (Requirement 3.1.3.9). The GPS outputs a National Marine Electronics Association (NMEA) Global Positioning System Fixed Data (GGA) string at a rate of five Hertz. Each NMEA string is stored onboard the payload. Also, the NMEA string is parsed for latitude, longitude, and mean sea level altitude. The parsed values are transmitted to the ground station.

The payload contains a liquid crystal display (LCD) screen. The LCD screen displays a color image. The LCD screen assists in the pre-launch checklist phase. The LCD screen displays sensor readings and the remaining power supply. The LCD screen mounts inside the payload such that an outside observer is able to clearly read the display. The LCD screen utilizes a dedicated power supply. The LCD screen is powered down a short time prior to launch.

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The payload includes a video camera. The video camera stores footage onboard the payload and is useful for educational outreach and public relations after launches. The video camera records both video and audio. The camera is an independent electrical system from the other payload electronics; therefore the video camera has a dedicated power supply. The video camera faces horizontally, and due to the clear acrylic payload housing, is able to record video from within the payload housing. Light emitting diodes ensures the camera is functioning properly prior to launch.

The official scoring altimeter is located within the payload section of the vehicle. The altimeter is mounted such that the scoring official can mark the device as the official altimeter (Requirement 1.2.2.1). The altimeter mounts in a secure manner to protect the sensor throughout the entire flight. The altimeter detects liftoff, records apogee, and reports the official altitude post-flight through a series of beeps (Requirement 1.2.1). The official altimeter is commercially available (Requirement 1.2). The altimeter is on a separate circuit from other payload electronics; therefore, the altimeter utilizes a dedicated power supply. In the original proposed design, the official altimeter is located in the same compartment as the parachute deployment altimeters. The decision to move the official altimeter to the payload is made based on the fact that when the black powder charges fire there is a chance that the spike in pressure could leak into the recovery altimeter compartment. This pressure spike jeopardizes the accuracy of the apogee altitude measurement. The clear payload housing also allows visual inspection of the official altimeter.

Required Subsystems

Atmospheric Data Gathering

Microcontroller

The atmospheric sensors send readings to an Arduino Mega 2560-R3 microcontroller, depicted in figure 32.

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Figure 32 Arduino Mega 2560-R3 Microcontroller

The microcontroller includes an ATmega2560 microprocessor which has 16 analog input ports, 14 digital pulse width modulation enabled output ports, and 54 general purpose digital input and output ports. The ATmega2560 microprocessor is capable of supporting at least four serial data interface devices, communicating with Inter-Integrated Circuit (I2C) data interface devices, and containing a Serial Peripheral Interface (SPI) bus for communicating with SPI data interface devices. The microcontroller has built-in 3.3 volt and five volt voltage regulators. The preceding characteristics are some of the determining factors to the microcontroller selection. The manufacturer’s data sheet for the Arduino 2560-R3 can be found in Appendix J.1.

Alternative microcontrollers are considerations. The major alternative to the Arduino Mega is the mbed NXP LPC1768. One desirable characteristic of the mbed microcontroller is the faster clock speed. The mbed incorporates a 100 megahertz clock speed. However, the mbed compiler is not as user friendly as the Arduino. Also, the open source nature of the Arduino allows for more example code and user advice. Another consideration is the Stellaris LM4F120 LaunchPad created by Texas Instruments. This microcontroller is relatively new. Again, the high clock speed is a desirable characteristic. The Stellaris Launchpad clock speed is 80 megahertz. Because the Stellaris Launchpad is a new product, back- orders and shipping time create a problem with using this product. The arrival date for this component is in mid-December, which is long after development begins. Table 38 lists the top four candidates for microcontroller selection. In addition to this information, the availability, size, power consumption, and cost are

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considerations throughout the selection process. The Arduino 2560-R3 is the microcontroller selection.

Clock Dimensions Input Current Part Name Distributor Speed (LxW) Mass Voltage Draw Cost 2.125" x 20 - Arduino 2560-R3 Sparkfun 16MHz 4.3125" 2.3oz 7 - 12V 200mA $58.95 mbed NXP 1 - 4.5V - 100 - LPC1768 Sparkfun 100MHz 1" x 2" 2oz 9V 200mA $59.95 1.06o Arduino Uno- R3 Sparkfun 16MHz 2.32" x 2.95" z 7 - 12V 50mA $29.95 Stellaris LM4F120 Texas 1 - 4.75 - 23 - LaunchPad Instruments 80MHz 2" x 2.25" 2oz 5.25V 300mA $4.99 Table 38 Top Four Candidates for Microcontroller Selection

Data Storage Device

The payload saves all sensor data to a 16 gigabyte high capacity micro secure digital (SD) card. The micro SD card organizes data into a 32 bit file allocation table (FAT32) file system. A micro SD card allows for quick retrieval of data. Another consideration for data storage is Electrically Erasable Programmable Read-Only Memory (EEPROM). However, retrieval of data from EEPROM is not as convenient. Also, EEPROM has a lower data storage capacity. Another consideration is the ENV-32X Embedded Data Logger. While the ENV-32X has a very user-friendly interface, it has a limited storage capacity. Finally, a standard size SD card is a possibility, but in this case the micro SD card’s smaller size is the determining factor. Table 39 lists the top four choices for data storage medium.

Part Name Part Number Distributor Capacity Cost micro SDHC 3FMUSD16FB-R Walmart 16 GByte $9.99 I2C EEPROM 525 Sparkfun 256 Kbit $1.95 ENV-32X 11196 Sparkfun 32 Mbit $34.95 SDHC 3FMSD32GBC10-R Walmart 32 GByte $17.99 Table 39 Top Four Choices for Data Storage Medium Upon selection of the micro SD card for data storage, a means of interfacing the micro SD card with the Arduino Mega 2560 is the next necessary determination in the design process. Several distributors manufacture breakout boards specifically designed to allow microcontrollers to easily communicate with micro SD cards over the SPI interface. The best-fit micro SD breakout board adapter is distributed by Adafruit. The board contains a built in five volt to three volt regulator. There are light-emitting diodes (LED) to indicate when the micro SD card is read from or

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written to. There is also a convenient locking mechanism to secure the micro SD card and allow for easy removal of the card. The board utilizes the SPI data interface to communicate with Arduino Mega 2560. Figure 33 shows the Adafruit micro SD adapter.

Figure 33 Adafruit Micro SD Adapter

Table 40 is a trade and selection table listing the top three candidates for micro SD adapter selection.

Part Data Dimensions Input Current Number Distributor Interface (LxW) Mass Voltage Draw Cost 254 Adafruit SPI 1.25" x 1" 0.12oz 3 - 5V 150mA (max) $15.00 32312 Parallax SPI 1.11" x 1" 0.11oz 3.3V 0.5mA $14.99 544 Sparkfun SPI 0.94" x 0.94" 0.09oz 5V 0.5mA $9.95 Table 40 Micro SD Adapter Selection Pressure In order to gather atmospheric pressure data, multiple barometric pressure sensors are being researched. For redundancy, two separate barometric pressure sensors are onboard the payload. For prototyping purposes, each sensor comes preinstalled on a breakout board. The primary selection is a BOSCH BMP180. The BMP180 uses a piezo-resistive sensor to detect applied pressure at a relative accuracy of plus or minus 0.017 pounds per square inch (psi). The piezo-resistive sensor outputs an analog value, which converts to a 16 bit digital value by way of an analog to digital converter. The BMP180 communicates with the Arduino Mega 2560 over the I2C data bus. The BMP180 data sheet can be found in Appendix J.2. Figure 34 is a picture of the BMP180 installed on a breakout board.

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Figure 34 BMP180 Breakout Board

The redundant barometric pressure sensor selection is the MS5611-01BA03 by Measurement Specialties. The MS5611-01BA03 also uses a piezo-resistive sensor. The relative accuracy is plus or minus 0.029 pounds per square inch. While the BMP180 utilizes a 16 bit analog to digital converter, the MS5611- 01BA03 implements a 24 bit analog to digital converter. It communicates with the Arduino Mega using the I2C data interface or the SPI data interface. The MS5611- 01BA03 datasheet is in Appendix J.3. Figure 35 is a picture of the MS5611- 01BA03 installed on a breakout board.

Figure 35 MS5611-01BA03 Breakout Board

Additional barometric pressure sensors are considered for selection. The determining characteristics are availability, data interface compatibility, precision of measurement, power consumption, and price. Table 41 is a pressure sensor trade and selection table.

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Part Data Precision of Input Current Name Manufacturer Distributor Interface Measurement Voltage Draw Cost DSS 1.8 - 32µA BMP180 BOSCH Circuits I2C ±0.017psi 3.6V (max) $15.00 MS5611- Measurement 1.8 - 12.5µA 01BA03 Specialties CSG Shop SPI - I2C ±0.029psi 3.6V (max) $29.99 Freescale MPL115A Semiconductor 2.375 - 10µA 1 Inc. CSG Shop SPI ±0.145psi 5.5V (max) $11.99 1.8 - 12µA BMP085 BOSCH Sparkfun I2C ±0.435psi 3.6V (max) $19.95 Table 41 Pressure Sensor Trade and Selection

Temperature

Temperature data is gathered by the same sensors gathering pressure data. The same piezo-resistive sensor is able to measure pressure and temperature. Therefore, because there are redundant barometric pressure sensors, there are also redundant temperature sensors onboard the payload. The purpose of this decision is to maximize efficiency of the payload electronics. In contrast, an individual sensor for the sole purpose of measuring temperature increases power consumption and decreases room for other electronics in the payload. Table 42 is a trade and selection table for the top considerations for temperature sensor selection.

Precision of Part Name Manufacturer Distributor Measurement Cost DSS $15.0 BMP180 BOSCH Circuits ±1.8° F 0 MS5611- Measurement $29.9 01BA03 Specialties CSG Shop ±1.44° F 9 TMP102 Texas Instruments Sparkfun ±1.8° F $5.95 $35.0 SHT11 Parallax Adafruit ±3.6° F 0 Table 42 Top Considerations for Temperature Sensor Selection

Relative Humidity

The payload consists of two redundant humidity sensors. The first selection is the Honeywell HIH4030. The HIH4030 is an analog sensor with no ADC. The analog output transmits to an analog input pin of the Arduino Mega 2560. The HIH4030 sensor uses a thermoset polymer capacitive sensing element to measure relative humidity to an accuracy of plus or minus 3.6 percent. For prototyping purposes, Sparkfun retails a sensor preinstalled on a breakout board that meets the

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necessary specifications. The sensor datasheet can be found in Appendix J.4. Figure 36 is a picture of the HIH4030 installed on a breakout board.

Figure 36 HIH4030 Breakout Board

The redundant relative humidity sensor is the HH10D Hope RF relative humidity sensor. In HH10D is also a capacitive type humidity sensor. The HH10D contains two humidity measurement gathering chambers. The two measurements save to EEPROM installed on the board. For precision purposes the two readings are averaged to obtain accurate data. The HH10D communicates with the Arduino Mega 2560 by means of a single digital frequency output. The sensor also utilizes the I2C data interface to relay an initial calibration value. The HH10D datasheet is in Appendix J.5. Figure 37 is a picture of the HH10D.

Figure 37 HH10D Breakout Board

In addition to the two chosen humidity sensors, alternative sensors are researched. Availability, data communication compatibility, data precision, power

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consumption, and price are all determining factors throughout the selection process. Table 43 is a trade and selection table of humidity sensors.

Part Data Precision of Input Current Name Manufacturer Distributor Interface Measurement Voltage Draw Cost HIH403 200 - 0 Honeywell Sparkfun Analog ±3.6% RH 5V 500μA $16.95 2.7 - HH10D Hope RF Sparkfun I2C ±3% RH 3.3V 150μA $9.95 2.4 - SHT11 Sensirion Adafruit Serial ±3% RH 5.5V 90μW $35.00 HIH613 2.3 - 1mA 0 Honeywell Sparkfun I2C ±4% RH 5.5V (max) $29.95 Table 43 Humidity Sensors Trade and Selection

Solar Irradiance

The device that measures solar irradiance is referred to as a pyranometer. The current payload design consists of redundant pyranometers. The payload consists of two of each chosen pyranometer. Each pyranometer has a field of view of only 180 degrees; therefore they are opposed 180 degrees to each other. The first pyranometer selection is the Apogee Instruments SP-110. The SP-110 measures the spectral light range from 380 to 1120 nanometers. The sensor outputs an analog value from zero to 350 millivolts. An increase of one millivolt corresponds to a radiation increase of 0.456 Watts per square foot. The analog output connects to an analog input pin of the Arduino Mega 2560. The SP-110 datasheet is located in Appendix J.6. Figure 38 is a picture of the SP-110.

Figure 38 SP-110

The chosen redundant pyranometer is a TAOS TSL2561. The TSL2561 measures the spectral range from 300 to 1100 nanometers. The sensor includes two photodiodes. The two photodiodes output an analog value which is then converts

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to digital by onboard ADCs. Therefore, in contrast to the SP-110, the TSL2561 outputs a 16 bit digital value. As stated in the TSL2561 datasheet, a digital output minimizes noise interference. The sensor communicates with the Arduino Mega 2560 over the I2C data interface bus. Through the use of the I2C interface, each individual photodiode is read separately. While the sensor’s formal use is an ambient light sensor, as stated in the datasheet, the raw output of the sensor is irradiance which is useful to calculate ambient light. The TAOS TSL2561 datasheet is located in Appendix J.7. Figure 39 is a picture of the sensor preinstalled on a breakout board.

Figure 39 TAOS TSL2561 Breakout Board

An alternative pyranometer is selected and will be tested to determine performance characteristics. Determining factors in the pyranometer selection process are availability, data communication compatibility, spectral range of measurement, power consumption, and price. Table 44 is a trade and selection table pyranometers.

Part Data Spectral Input Current Dimensions Name Distributor Interface Range Voltage Draw (LxWxH) Cost Apogee 380 - 0.94" x 0.94" SP-110 Instruments Analog 1120nm 0V 0V x 1.08" $169.00 300 - TSL2561 Adafruit I2C 1100nm 2.7 - 3.6V 0.5mA 0.75" x 0.75" $12.50 TEMT600 360 - 0.375" x 0 Sparkfun Analog 970nm 6V 20mA 0.375" $4.95 Table 44 Pyranometers Trade and Selection

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Ultraviolent Radiation

Ultraviolet (UV) radiation is the light spectrum between ten nanometers and 400 nanometers. This range of the light spectrum is separated into three categories, UVA, UVB, and UVC. The payload measures these three categories which include the range of the light spectrum from 210 nanometers to 400 nanometers. When selecting a sensor to gather UV radiation data, a deciding characteristic is the range of the light spectrum each sensor measures. The payload includes redundant UV sensors that measure the total UVA, UVB, and UVC spectrums. Each sensor has a field of view of 180 degrees; therefore it is necessary to place the sensors opposing 180 degrees to each other. The first UV sensor selection is the Apogee Instruments SU-100. The SU-100 measures the light spectrum from 250 nanometers to 400 nanometers. A protective dome houses the sensor. The SU-100 requires no voltage source, but rather accumulates power from the sun. The SU-100 datasheet is in Appendix J.8. Figure 40 is a picture of the SU-100.

Figure 40 SU-100

The redundant UV sensor selection is the sglux TOCON_ABC3. The TOCON_ABC3 is a preamplified UV sensor. The sensor measures the UV light spectrum from 210 nanometers to 380 nanometers. The sensor outputs an analog voltage representing UV light measurements. The analog output connects to an analog input of the Arduino Mega 2560. In contrast to the SU-100, the TOCON_ABC3 requires an input voltage of at least five volts. The TOCON_ABC3 datasheet is in Appendix J.9. Figure 41 is a picture of the TOCON_ABC3.

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Figure 41 TOCON_ABC3

Alternative UV sensors are products that are past considerations. The main alternative to the two chosen sensors is the Solar Light Inc. PMA1107. A negative feature of the PMA1107 is the high price. When choosing UV sensors, the determining characteristics are availability, spectral measurement range, power consumption, size, and price. Table 45 is a trade and selection table of the three considerations for UV sensor.

Spectral Input Current Dimensions Part Name Distributor Range Voltage Draw (LxWxH) Cost Apogee 250 - 0.925" x 0.925" SU-100 Instruments 400nm 0V 0A x 1.08" $159.00 TOCON_ABC 210 - 3 sglux 380nm 2.5 - 15V 0.8mA 0.4" x 0.4" x 0.3" $148.00 Solar Light 260 - PMA1107 Inc. 400nm 5 - 12V 1mA 1.6" x 1.6" x 1.8" $525.00 Table 45 UV Sensor Trade and Selection

Global Positioning System

The chosen GPS is the Locosys LS20031. The LS20031 contains a GPS antenna, a MC-1513 GPS module, and a transistor-transistor logic (TTL) data interface. The GPS has a refresh rate range of one Hertz to ten Hertz. The initial testing refresh rate is five Hertz. Through testing, the optimal refresh rate is to be determined. The GPS is also capable of receiving data from 66 different channels. Again, the optimal channel is to be determined through testing. The GPS is accurate to within nine feet. The GPS receives a NMEA string. There are six possibilities for NMEA output message selection: Global positioning system fixed data (GGA), geographic position – latitude/longitude (GLL), GNSS and active satellites (GSA), GNSS satellites in view (GSV), recommended minimum specific GNSS data (RMC), and course over ground and ground speed (VTG). The research concerning the GGA output message which contains the following data: coordinate universal time, latitude, longitude, number of satellites used, horizontal dilution of precision, mean

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sea level altitude, and geoid separation. The LS20031 datasheet is in Appendix J.10. Figure 42 is a picture of the LS20031.

Figure 42 LS20031

Alternative GPS modules are available to conduct research. Although other GPS modules have similar characteristics, the LS20031 is the best-fit for the project; because, multiple members of the team have experience working with the module. Also, availability, data communication compatibility, refresh rate, power consumption, and price are determining factors for the decision. Table 46 is a trade and selection table of the top three GPS modules that are considerations for this type of a project.

Refresh Dimensions Input Current Part Name Distributor Interface Rate Mass (LxWxH) Voltage Draw Cost 0.49o LS20031 Locosys Serial 1- 10Hz z 1.18" x 1.18" 3 - 4.2V 29mA $60.00 1.0" x 1.35" MTK3339 Adafruit Serial 1 - 10Hz 0.3oz x 0.25" 3 - 5.5V 25mA $39.95 UBlOX I2C, 0.33o 100mA MAX-6Q CSG Shop Serial 5Hz z 1.77" x 0.59" 3.3V (max) $69.99 Table 46 GPS Modules Trade and Selection

Wireless Transmitter

The wireless transmitter that relays atmospheric sensor readings and GPS data is the XBee-PRO XSC S3B. The XBee-PRO XSC S3B is capable of transmitting telemetry to a line of sight range of 28 miles when coupled with a high gain antenna. The transmitter operates at the 900 Megahertz frequency band. The 900 Megahertz band splits into 12 different channels to help prevent interference from other devices transmitting at this frequency. The programming of the transmitter incorporates a 64 bit personal area network identification number. The XBee-PRO XSC S3B requires a low voltage of only 2.4 volts to 3.6 volts. The XBee-PRO XSC

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S3B datasheet is available in Appendix J.11. Figure 43 is a picture of the XBee- PRO XSC S3B.

Figure 43 XBee-PRO XSC S3B

The XBee brand is the only major consideration for wireless transmission in the group’s view; however, XBee offers products that transmit over two different frequency bands. The chosen transmitter utilizes the 900 Megahertz band while the alternative XBee consideration utilizes the 2.4 Gigahertz band. The 2.4 Gigahertz XBee has a significantly lower telemetry range; although the 900 Megahertz XBee has a lower data transmission rate. The 2.4 GHz XBee data transmission rate is 250 kilobits per second and the 900 Megahertz XBee data transmission rate is only 10-20 kilobits per second. The telemetry range outweighs the data transmission rate in terms of the team’s priorities. Table 47 is a trade and selection table consisting of the two considered XBee modules.

Telemetry Frequency Data Input Current Part Name Distributor Range Band Rate Voltage Draw Cost XBee-PRO 10 - 215mA XSC S3B Digi 28 miles 900MHz 20Kbps 2.4 - 3.6V (max) $42.00 XBee-PRO 205 mA ZB Digi 2 miles 2.4GHz 250Kbps 2.7 - 3.6V (max) $28.00 Table 47 XBee Module Trade and Selection

Autonomous Camera Orientation System

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The autonomous camera orientation system consists of a microcontroller, photographic camera, accelerometer, data storage device, servo motors, and a dedicated power supply. The camera continuously takes pictures throughout the entire flight and saves them to the data storage device. Upon deployment of the main parachute, the autonomous camera orientation system activates. The accelerometer relays attitude measurements of pitch and yaw to the microcontroller. The microcontroller then calculates the appropriate correction angles for two servo motors. The camera mounts to the servos. Five minutes after landing, the autonomous camera orientation system deactivates.

Microcontroller

The microcontroller that determines the servo motors’ angle is the Arduino Pro Mini 328. The Arduino Pro Mini 328 contains an ATmega328 microprocessor. The ATmega328 has significantly fewer digital and analog ports then the ATmega2560. Fewer ports allow for smaller overall size. The Arduino Pro Mini 328 is only 0.7 inches by 1.3 inches, and has a mass of only 0.07 ounces. The ATmega328 is capable of communicating with devices using the SPI data interface, I2C data interface, and the serial data interface. The Arduino Pro Mini 328 datasheet can is accessible in Appendix J.12. Figure 44 is a picture of the microcontroller.

Figure 44 Arduino Pro Mini 328 Mircocontroller

Other considerations for alternative microcontrollers did not meet the standards the team looked for. The mbed NXP LPC11U24 is the secondary consideration. The higher clock speed was a desirable characteristic, but the mbed compiler is a negative factor throughout the mbed NXP LPC11U24’s consideration. The compiler is only accessible when Internet access is available. This constraint seems unacceptable. Microcontroller selection centers on features including availability, clock speed, size, power requirements, and price. Table 48 is a trade and selection table of the seriously considered microcontrollers.

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Clock Dimensions Input Current Part Name Distributor Speed (LxW) Mass Voltage Draw Cost Arduino Pro Mini 328 Sparkfun 16MHz 0.7" x 1.3" 0.07oz 5 - 12V 150mA $18.95 mbed NXP 2.4 - LPC11U24 Sparkfun 48MHz 2.13" x 1.02" 1 - 2oz 3.3V 50mA $44.95 Stellaris LM4F120 Texas 4.75 - 23 - LaunchPad Instruments 80MHz 2" x 2.25" 1 - 2oz 5.25V 300mA $4.99 Table 48 Microcontrollers Trade and Selection

Camera

The selected photographic camera is the VC0706. Adafruit distributes a breakout board with a preinstalled module. The camera sends each photograph to the Arduino Pro Mini. The separate microcontroller is necessary due to the expected delay time when saving the picture files to the micro SD. The VC0706 communicates with the Arduino using the serial data interface bus. The photographs save to the micro SD card in Joint Photographic Experts Group (JPEG) format. The camera mounts from the servo motors onto the breakout board using the prefabricated mounting holes. The VC0706 datasheet can is accessible in Appendix J.13. Figure 45 is a picture of the camera.

Figure 45 VC0706 Photographic Camera

Alternative camera selections are past considerations. The Sparkfun SEN-10061 is the first to attain status as a serious alternative. The major negative characteristic of the SEN-10061 is the lower resolution of the pictures. There are multiple figures available for the resolution of the SEN-10061. The distributor website, Sparkfun.com, states that resolution is 640 pixels by 480 pixels. Table 49 is a trade and selection table contacting the top three seriously considered camera modules.

Part Distributor Data Pixels Dimensions Input Current Cost

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Name Interface (LxWxH) Voltage Draw 1.26" x 1.26" x VC0706 Adafruit Serial 640 x 480 0.25" 5V 75mA $42.00 SEN- 1.26" x 1.26" x 10061 Sparkfun Serial 160 x 120 0.25" 5V 0 - 100mA $49.95 Electro 0.24" x 0.24" x OV7670 Dragon I2C 640 x 480 0.16" 3.3V 60mW $9.30 Table 49 Camera Modules Trade and Selection

Servos

Servos control the orientation of the camera. One servo mounts to the aluminum rail on the side of payload. The second servo mounts to the first servo’s rotational arm. The camera mounts to the second servo’s rotational arm. This setup allows for camera orientation in the horizontal and vertical directions. There are many types of servo motors. The optimal servo for the payload has characteristics of being small but not so small that the torque compromises the function as a result. Both servo motors are HS-85BB+ Mighty Micros. The Mighty Micro is small, lightweight, and strong. A pulse width modulated output signal from the Arduino Pro Mini 328 controls the servo. The gears within the Mighty Micro are nylon. Another possible material for gears is metal. Servo motors that contain metal gears burn out if they become stuck. In contrast, when servo motors contain nylon gears, usually the gears strip before the motor burns out. The Mighty Micro also contains a ball bearing to make the motor move smoother. The HS-85BB+ Mighty Micro datasheet is accessible in Appendix J.14. Figure 46 is a dimensional layout of the servo.

Figure 46 Servo Dimensional Layout

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Alternative servo motors are prior considerations of the team. The Standard Servo f by Parallax is definitively too large. The HS-55 Sub-Micro is too small and weak. Table 50 is a trade and selection table of the top three seriously considered servo motors. The determining characteristics are availability, torque, size, power requirements, and price.

Input Current Part Name Distributor Torque Mass Dimensions Voltage Draw Cost HS-85BB+ 41.66 1.14" x 0.51" x Mighty Micro ServoCity oz-in 0.67oz 1.18" 4.8V 240mA $19.99 Standard Servo 2.2" x 0.8" x (#900-0005) Parallax 38oz-in 1.55oz 1.6" 6V 190mA $12.99 HS-55 Sub- 15.27 0.89" x 0.45" x Micro ServoCity oz-in 0.28oz 0.94" 4.8V 150mA $9.99 Table 50 Servo Motors Trade and Selection

Accelerometer

An accelerometer measures acceleration in multiple planes. Because gravity is essentially constant, it is possible to accurately measure the tilt by relative acceleration. The chosen accelerometer is an ADXL345. The ADXL345 measures acceleration in the X, Y, and Z planes. The ADXL345 utilizes the I2C data interface bus. The X and Y measurements of the accelerometer represent the tilt and yaw of the payload. Each reading stores to the micro SD card for post flight analysis. The measurements also determine the angle of the servo motors. The ADXL345 datasheet is accessible in Appendix J.15. Figure 47 is a picture of the sensor following installment on a breakout board.

Figure 47 ADXL345 Breakout Board

Alternative accelerometers are in consideration for the project. The most serious alternative accelerometer is the Memsic Mx2125. The sensor is still in consideration. The Mx2125 utilizes a pulse width modulated output for each of the axis measurements. Both the Mx2125 and the ADXL345 are facing head to head

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testing by the Tarleton Aeronautical Team. Through testing, the most reliable sensor remains to win its spot. The Mx2125 is only capable of measuring acceleration of plus or minus 3Gs, and the ADXL345 is capable of measuring plus or minus 16 Gs. This is the determining factor for the current selection. However, the ADXL345 measures three planes of acceleration while the Mx2125 measures only two. Two planes of measurement are all that is necessary for the camera orientation system. Table 51 is a trade and selection table of the serious competition among accelerometers.

Part Range of Input Current Name Distributor Interface Axes Measurement Voltage Draw Cost ADXL345 Sparkfun I2C 3 ±16 g 2 - 3.6V 40μA $27.95 Mx2125 Parallax PWM 2 ±3 g 3.3 - 5V 5 mA $29.99 ADXL335 Sparkfun I2C 3 ±3 g 1.8 - 3.6V 320μA $24.95 Table 51 Accelerometers Trade and Selection

Video Capture

Video Camera

The payload houses a video camera. The chosen video camera is the VCC-003- MUVI-BLK. The camera is extremely small, lightweight, and capable of recording video for up to 90 minutes. The camera has a rechargeable battery preinstalled. The camera utilizes a dedicated micro SD card. The selected micro SD card that stores the video is the same type of micro SD selected to save sensor data. The camera powers on manually very near to launch time. The camera is completely independent from the other payload electronics. The camera mounts to the rails of the payload. The resolution of the video is 640 pixels by 480 pixels. After the flight, the micro SD card removes manually from the camera and the footage facilitates educational engagement purposes and public relations. Table 52 is a datasheet of the VCC-003-MUVI-BLK. Figure 48 is a picture of the selected video camera.

Figure 48 VCC-003-MUVI-BLK Video Camera

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Research on alternative video cameras led to this selection. The most competitive video camera is the FlyCamOne2. The FlyCamOne2 has the same capacity for length of video. However, all distributors of FlyCamOne2 state that the product not currently available. Table 50 is a trade and selection table of the top two competing video cameras.

Frame Dimensions Battery Usage Part Name Distributor Rate Mass (LxWxH) Time Cost VCC-003- 1.10" x 2.17" x MUVI-BLK Adorama 20 - 25fps 1.76oz 0.79" 70 - 90 minutes $49.95 1.57" x 3.15" x FlyCamOne2 Hacktronics 25fps 1.31oz 0.55" 30 minutes $69.95 Table 52 Top Two Video Cameras Trade and Selection

Liquid Crystal Display

Sparkfun distributes the selected LCD, part number 11062, which mounts on a breakout board. The LCD has a color display with a resolution of 132 pixels by 132 pixels. The Arduino Mega 2560 microcontroller controls the LCD. The LCD screen utilizes the SPI data interface bus. One desirable feature of the chosen LCD screen is the low voltage requirements. At a minimum, the LCD screen requires 3.3 volts. The LCD-11062 datasheet is available in Appendix J.17. Figure 49 is a picture of the chosen LCD screen.

Figure 49 Sparkfun LCD-11062 Screen

Alternative LCD screens are serious past considerations. A smaller black and white LCD and a larger touch screen color LCD are the two most seriously competitive alternatives. The black and white LCD falls short of some needed conventions; because, it is a desirable feature to have color rather than black and white. The touch screen falls short; because, the functionality is not necessary.

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Table 53 is a trade and selection table of the three most serious competitors for LCD screens.

Part Dimension Input Number Distributor Resolution Color s Voltage Current Draw Cost LCD- 132 x 132 11062 Sparkfun pixels Yes 1.5" x 2.5" 3.3 - 6V 108 – 324mA $34.95 LCD- 48 x 84 1.75" x 10168 Sparkfun pixels No 1.75" 3.3 - 6V 240 – 320mA $9.95 240 x 320 335 Adafruit pixels Yes 2.5" x 3.2" 3.3 - 5V 150mA $40.00 Table 53 LCD Screen Trade and Selection

Official Scoring Altimeter

The official scoring altimeter beeps the apogee height upon landing. The height records in feet in accordance with the altimeter’s data. The chosen official scoring altimeter is the Adept A1E. The altimeter wiring is completely independent of the other payload electronics; therefore, the altimeter utilizes a dedicated power supply. The required power supply is a twelve volt battery which comes with the altimeter upon the receipt of the order. The battery that ships with the altimeter is the GP-23A Alkaline Lighter Battery. This power supply allows the altimeter to function for up to 10 hours. The altimeter datasheet is available in Appendix J.18. Figure 50 is a picture of the Adept A1E.

Figure 50 Adept A1E

The Adept A1E is the strongest altimeter due to its small size and long battery life. Alternative scoring altimeters offer strong competition. The alternatives are able to fire black powder and record the apogee altitude. The PerfectFlite Stratologger is one of the altimeters recovery deployment employs. The decision to select an altimeter that is specifically designed for only reading out the apogee altitude in beeps and not a multifunctional deployment altimeter fits the parameters best. Similar reasoning was leads to the rejection of the alternative official scoring

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altimeter. Table 54 is a trade and selection table of the three serious considerations for official scoring altimeters.

Name Distributor Dimensions Input Voltage Cost A1E Adept 0.55" x 2.2" 12V $29.95 Stratologger PerfectFlite 2.75" x 0.9" 9V $79.95 ARTS Loki Research 3.75" x 1.4” 9 - 15V $190.00 Table 54 Official Scoring Altimeters Trade and Selection

Power Supply

Eight nine volt batteries power the payload. This power supply provides more than adequate voltage to power all the electronic components. Voltage regulators level the voltage down to five volts and three volts, depending on requirements for each component. The payload is able to function at full operation for at least two hours. The power supply distributes power to the components so that if single batteries fail the entire system will not fail.

Battery

Research and analysis lead to the battery selection in place. The determining characteristic is capacity. The U9VLBP has a large capacity of 1.2 amp-hours. This capacity exceeds the current requirements of the payload. Alternative batteries are prior considerations, but both alternatives have a lower capacity. However, both alternatives are more readily available. Table 55 is a trade and selection table of the three most serious considerations for batteries.

Part Name Manufacturer Capacity Voltage Price U9VLBP Ultralife 1.2 Ah 9V $5.50 Coppertop Alkaline Duracell 580 mAh 9V $11.00 Advanced Lithium Energizer 750 mAh 9V $9.49 Table 55 Batteries Trade and Selection

Performance Characteristics

The performance of the payload system characterizes how well the system completes the necessary objectives. The main objectives of the payload are to store and transmit atmospheric and GPS data, to collect pictures at the proper

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orientation, and to record flight footage. Research and analysis of many sensors and electrical components view evaluation criteria for the success of the payload system. Wiring diagrams and power budgets formulate and verify through analysis and testing. Researching cameras and formulating and prototyping mechanisms for leveling the camera throughout flight is a job necessary to sound design. Exact dimensional CAD drawings and 3D renderings illustrate and help evaluate the fitment of the camera orientation system. Verification of performance through prototyping and flight testing is ongoing. Evaluations of video cameras and their ability to record video footage are paramount for the objectives of the flight. Testing the video camera lends itself to verification of the camera. Table 56 represents the evaluation and verification metrics for the payload subsystems.

Performance Subsystem Evaluation Verification Characteristics

Research and Analysis of schematics Atmospheric Data Store and transmit analysis of and testing of Gathering atmospheric and GPS data sensors and components electronics

Save pictures of the proper Research and Prototyping and flight Camera Orientation orientation CAD drawings testing Store footage of the entire Research of video Video Footage Testing flight cameras Table 56 Payload Subsystems Evaluation and Verification Metrics

Verification Plan

The verification plan reflects how each requirement to the payload system is in a satisfactory state of completion. The SOW enumerates requirements that appear in brevity in Table 57. The table also lists satisfying features of the design for the requirement in question. Ultimately, design features require verification to ensure that it actually meets the requirement. This goal is achievable through testing, analysis, and/or inspection.

SOW Satisfying Design # Requirement Feature Verification Method

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Launch vehicle shall carry 3.1 a science or engineering SMD payload selection Inspection, Analysis payload

Measurements of An appropriate sensor pressure, temperature, for each of these Testing, Analysis, 3.1.3.1 relative humidity, solar measurements has been Inspection irradiance and ultraviolet selected radiation

Measurements occur Main flight 3.1.3.2 every 5 seconds during computer/microcontroller Testing, Analysis descent sampling rate

Main flight Measurements occur 3.1.3.3 computer/microcontroller Testing, Analysis every minute after landing sampling rate

Main flight computer/microcontroller Data collection will cease Testing, Analysis, 3.1.3.4 shall halt all data 10 minutes after landing Inspection collection after 10 minutes

Payload will take 2 Control of camera to 3.1.3.5 pictures during descent appropriate image Testing and 3 after landing capture

Pictures taken must be properly oriented such Self-leveling camera 3.1.3.6 that the sky is at the top Testing, Inspection mount of the frame, ground at the bottom

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microSD onboard Data from payload will be storage; radio link 3.1.3.7 stored onboard and between Xbee radios Testing transmitted wirelessly from rocket vehicle to ground station

Separations of payload at apogee is allowed, but The payload does not 3.1.3.8 such separation may detach from the rocket Inspection cause drifting outside vehicle recovery area

Locosys LS20031 is 3.1.3.9 Payload will carry GPS incorporated in the Inspection, Testing payload design

Data from payload will be (satisfying design collected, analyzed, and feature) experimental 3.2 reported by team Analysis logic, approach, and following scientific method of investigation method

Payload will be housed Payload must be in the main vehicle and 3.5 designed to be Testing, Inspection remain there during the recoverable and reusable entirety of the flight

Table 57 - SOW Verification

Preliminary Integration Plan

Payload Integration Plan

The payload integrates into the payload housing structure of the vehicle in a simple fashion. A payload framework consisting of two bulkheads connected to each other by two aluminum rails houses the payload. Since the payload housing structure is part of the vehicle, the payload framework uses dimensions to ensure compatibility between the payload and the vehicle.

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Preparing the payload for flight is simple prior to payload framework integration with the vehicle. Once the payload is ready for flight, the payload framework is ready for integration. At this time, the framework inserts from the lower portion of the payload housing structure. Once the payload framework is in place, it secures with the use of two screws from the exterior of the payload housing structure into the bulkhead. At this point, the physical integration of payload is complete.

Precision of instrumentation, repeatability of measurement, and recovery system.

The instrumentation that takes measurements is commercially available on breakout boards that manufacturers verify and test. Data sheets that manufacturers provide for each sensor outline the precision and error tolerance. Collected data compare to known data from sources such as NOAA to verify accuracy and precision. The repeatability is easily accomplishable, since the payload is modular and sensor data records to microSD cards. The only thing necessary to gather a new set of data is the removal of the payload to recover the data, install a fresh microSD card, and reload the payload into a fresh preparation of the launch vehicle for a new flight. The parachutes consist of durable, silicon- coated ripstop nylon. This allows the parachutes to function repeatedly, requiring only new black powder charges before preparing the recovery system for a new flight. Table 58 lists the precision of each sensor in the payload.

Purpose Product Precision Barometric Pressure BMP180 ±0.017psi Barometric Pressure MS5611-01BA03 ±0.029psi Temperature BMP180 ±1.8° F Temperature MS5611-01BA03 ±1.44° F Humidity HIH4030 ±3.6% RH Humidity HH10D ±3% RH Solar Irradiance SP-110 ±5% Solar Irradiance TSL2561 ±5% Ultraviolent Radiation SU-100 ±10%

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Ultraviolent Radiation TOCON_ABC3 ±10% GPS LS20031 ±9.84ft Accelerometer ADXL345 ±4.3mg Official Altimeter Adept A1E ±1ft Table 58 Payload Sensors Precision

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Drawings and Electrical Schematics

Figure 51 is a conceptual wiring diagram of the atmospheric data gathering subsystem. Determination of proper wiring was comes from a careful review of the datasheets.

Figure 51 Atmospheric Data Gathering Conceptual Wiring

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Figure 52 Camera Orientation Wiring Electrical Schematic

Figure 53 represents the schematic for the payload sensors. The schematic is fully developed to allow for the design and development of printed circuit boards (PCB). Though the current payload design uses an array of breakout boards, a PCB design would lead to a more compact and efficient payload system. Due to resolution limitations and the large size of the perf board circuit, the figure cannot properly display the labels of each component. To remedy this, a PDF is posted on the Tarleton Aeronautical Team website to show the schematic in full resolution.

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Figure 53 PCB Schematic for Payload Sensors

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Figure 54 I^2C Sensors

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Figure 55 Radio Communications

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Figure 56 MicroSD Card Reader

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Figure 57 Analog Sensors

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Figure 58 USB Interface

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Figure 59 GPS Module

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Figure 60 Voltage Regulators

PCB Layout

During the course of testing and design, the payload circuitry progresses towards a complete PCB layout. Figure 61 demonstrates the preliminary design for the PCB layout. A number of advantages arise from using and creating a custom PCB design. The size of the payload housing reduces drastically. Altogether the breakout boards add to almost 20 in2. The design uses a 4.75” x 4.75” surface and even then leaves a large amount of unused space and room for optimization. After thorough planning, the PCB design can reasonably be reduced to nearly half the size of a breakout board system.

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Figure 61 Preliminary PCB Layout

Furthermore, utilizing a custom-designed PCB layout optimizes the efficiency to the designer’s exact specifications. Figure 62 displays a component listing printout from the PCB design. Each component is specifically chosen to minimize the amount of ESR introduced to the system. Copper wires and through-hole components inherently have larger amounts of ESR. This causes parasitic capacitance and resistances to effect the circuitry adversely. Each component in figure is chosen for its SMT design. This smaller area of conductive material

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reduces the ESR drastically (from approximately 7Ω to 0.05Ω). The difference between these two component constructions is available in related figures.

Figure 62 Component Listing

Figure 63 SMT (left) vs. Through Hole Devices (right)

Ultimately, the PCB design will be optimized over the course of the testing phase to yield a fully-functional and extremely efficient payload circuit. The difficulty of this PCB design and build process sets this payload design apart from others. To create this preliminary design cost approximately 100 man-hours. To improve upon this design is a tedious and challenging process; however, it results in an exceptional payload circuit.

Cross-Component Compatibility

The atmospheric data gathering subsystem has a unidirectional data flow. The flow of data travels from the atmospheric sensors relaying measurements to the Arduino Mega 2560. Once sensor measurements come in, they transfer and save to a file on the 16 gigabyte micro-SD card. Pre-flight measurements display on the LCD screen just prior to launch throughout the flight. The Arduino Mega 2560 sends telemetry data to the XBee Pro XSC S2B at regular intervals. The XBee Pro

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XSC S2B wirelessly transmits the telemetry data to the ground station. Figure 54 portrays the flow of data for the atmospheric data gathering subsystem.

Figure 64 Atmospheric Data Gathering Subsystem Data Flow

The autonomous camera orientation subsystem also has a unidirectional flow of data. The data flow begins with the ADXL345 accelerometer measuring the tilt and yaw of the payload. The tilt and yaw data transfers to the Arduino Pro Mini. At regular intervals, the VC0706 photographic camera takes pictures and transmits them to the Arduino Pro Mini. The Arduino Pro Mini sends angular settings to the HS85BB Mighty Micro servo motors in accordance with the accelerometer readings while simultaneously saving pictures to the micro-SD card. Figure 55 models the data flow for the autonomous camera orientation system.

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Figure 65 Autonomous Camera Orientation System Data Flow

In order for the Arduino microcontrollers to properly function, the user must design and write software that governs the desirable function. The software dictates a protocol determining how each separate component communicates with the microcontroller. Software design and analysis follows the object oriented programming paradigm, allowing the code to easily update and debug. The software development process begins with conceptual designs. Implementation and deployment of code is the final phase of the process. This method generates efficient and redundant software. The software development process is under way. The first steps include the generation of flow charts, which function as high level layouts for the software. There are two programmable microcontrollers onboard the payload; therefore, two separate flow charts exist. Figure 66 is the flow chart for the atmospheric data gathering subsystem and Figure 67 is the flow chart for the autonomous camera orientation subsystem.

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Figure 66 Atmospheric Data Gathering Software Flow Chart

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Figure 67 Autonomous Camera Orientation Software Flow Chart

Payload Concept Features and Definition

Creativity and Originality

A clear acrylic section of the vehicle body houses the payload. This allows visual inspection of all sensors and indicator LEDs along with verification readouts from the LCD during assembly and pre-launch operations. Given the payload

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integration design, the payload is easily and quickly removable. Given the modular design of the power circuits, disposable batteries are replaceable when necessary.

Although testing and prototyping begins with breakout boards, ultimately the goal is to work towards designing the PCBs from scratch to size and wiring specifications. These completely original PCBs and the designs are realized as a final product. A silkscreen image prints a team (or university) emblem on the board in Tarleton purple.

Figure 68 - PCB Board

The self-leveling camera system ensures that photographs are taken in proper orientation as required. The tilt sensor detects any changes in orientation, where two servo motors correct movement. The LCD displays appropriate data and relevant readouts to ensure and verify payload functionality. The team plans to use video camera for flight documentation and plans to use this as an educational outreach tool.

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Figure 69 - Self-Leveling Camera System

Uniqueness and Significance The scientific payload is significant in that it meets the SMD requirements set forth in the NASA SLP SOW. All design features and component selections reflect meeting the customer prescribed specifications. The ability of the payload to gather atmospheric data is significant for the analysis of changing conditions as the vehicle varies in altitude. All data compiling is useful for finding correlations to altitude such that atmospheric conditions from ground level to vehicle apogee can model and test for significance. The clear acrylic airframe houses the payload. This is unique in that it allows visual inspection of all components of the payload externally. Once preparation and integration of the payload takes place in the vehicle, a final verification ensures flight readiness relating to the payload. An LCD screen displays relevant checks and indicators from the flight software. The means to maintain proper orientation of the camera is unique as well. Because the orientation of the payload is no guarantee upon descent, a corrective measure uniquely takes place. The chosen solution creates and implements a self- leveling camera. Any changes to the orientation of the payload relative to starting position at the launch pad can result in measurement and compensation for the discrepancy. A video camera in the payload captures video of vehicle flights. Although this is a standard practice in rocketry, video recordings provide flight documentation that can be useful in educational engagement and community outreach.

Suitable Level of Challenge The design complexity and implementation of a functional SMD payload is an extremely taxing challenge. Not only must appropriate components reside in the

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vehicle, they must orchestrate in a manner that achieves a useful atmospheric measuring instrument. This involves proper interfacing, sufficient power supply, and adequate programming logic of all constituent pieces of the payload. Although very challenging, design of the electrical circuits ultimately seek PCB implementation. This is a significant advantage over breakout board and perforated board mounting. The space saving and overall efficiency of the PCB design merits it as the fruition of majors aims of the payload electronics. Despite the level of challenge and time necessary to complete this design, it is worthwhile in view of the integrity of the payload system as a whole.

Science Value

Payload Objectives

The main objectives of the payload are to store and transmit atmospheric and GPS data, capture photos of the correct orientation, and to record video footage of the entire flight. The payload must receive sufficient power to operate while on the launch pad, during flight, and after landing. The atmospheric data collection illuminates the change in atmospheric variables depending on altitude. The GPS data shows the flight path of the vehicle and aids in recovery. The camera orientation system tests a multi-servo, autonomous orientation device. The video footage is profitable for public outreach in the aftermath of the flight.

Payload Success Criteria

If the payload objectives succeed, the mission is a success. The atmospheric data is in this case accurate and stores at a minimum of one reading from each sensor every five seconds. Transmission of data from the payload to the ground station operates throughout the entire flight. Two photos during descent and three after landing capture valuable references in the flight. The photos should orient in such that they portray the sky at the top of the frame and the ground towards the bottom. The video footage spans the entire flight. Table 59 summarizes the payload objectives, their science value and success criteria.

Payload Objective Science Value Success Criteria Represents Change in Store and Transmit Accurate Data, Collected at Required Atmospheric Variables Atmospheric and GPS Data Frequency, Transmitted Throughout Flight Depending on Altitude

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Test Multi-Servo Two Pictures During Descent and Three Camera Orientation Orientation Device after Landing, Correct Orientation Record Entire Flight, Collect Usable Record Video Footage Aid in Public Outreach Footage Table 59 Payload Objectives Summary

Experimental Logic, Approach, and Method of Investigation

The SMD payload gathers data from approximately 5280 feet above ground level (AGL) down to the landing site, collecting altitude-varied data for five atmospheric variables: pressure, temperature, relative humidity, solar irradiance, and ultraviolet radiation. This data determines the accuracy of the payload sensors and the statistical correlations between each of the various variables. Together, these two calculations develop a regression model for each variable. By creating a model to represent these correlation effects, a new and comprehensive formula could demonstrate causal relationships between these five variables or any derivative subset.

Regression Model

With a large number of samples ranging across the various test and demonstration flights, the data plots determine the order of polynomial of best fit. If the order of this polynomial is one then a simple linear regression model of the form can result through statistical analysis of the data. If the order of the proposed polynomial does not equal one, then local regression determines the most likely model to represent the data and predict new values. This process is repeatable for each variable against altitude and then in any combination of subject variables. Once this model establishes itself and verifies, it can stand to test and improve any functions relating altitude to any of the variables this study includes.

Correlations between SMD Sensor Readings

After establishing an acceptable model, the distributions of each variable aid in finding covariance between any two variables and, consequently, the correlation between those variables. After determining the correlation, the model can evolve as necessary to provide more accuracy.

Accuracy of Sensors

Though the datasheets for each sensor propose a certain level of accuracy, this cannot guarantee the sensors will perform to this level within the payload circuitry. By comparing collected data against atmospheric measurements from national databases, any discrepancy can establish itself. Furthermore, using both data sets can establish confidence intervals to ensure new data are within a particular range

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of the presumably absolute readings from these databases. This gives a measure of accuracy as it pertains specifically to the payload circuitry.

Test and Measurement, Variables, and Controls

The variables include altitude, temperature, atmospheric pressure, relative humidity, solar irradiance, UV radiation and random electrical noise in the payload circuitry. The -terms in the example model above comprise of vectors of each individual payload measurement. The –term’s resultant vector represents the unobservable electrical noise in the system.

The controls in consideration are the measurements from national agencies, such as the National Oceanic and Atmospheric Administration (NOAA), whose measurement devices have presumably negligible errors. The measurement process centers on data acquisition from the payload sensors. The testing process occurs post-flight in the form of a statistical analysis of the acquired data.

Relevance of Expected Data and Accuracy/Error Analysis

The findings compare against documented formulae for atmospheric measurements to determine their validity. Furthermore, the models could potentially represent undocumented relations between these variables. The accuracy and error analysis are an inherent part of the entire process. All errors factor into the final regression model. Small differences from the expectant values will demonstrate as random noise in the model.

Preliminary Experiment Process Procedures

Numerous hypotheses exist prior to measurement and testing, but the testing of these hypotheses still requires the analysis of future data. Thus, no actions in experimentation are complete.

Safety and Environment (Payload)

REFER TO SAFTEY AND ENVIROMENT (VEHICLE) V) Project Plan

Budget Plan

A summary of the costs for various aspects of the project assembles into a chart format for efficiency and ease of use. Tables of pertinent data follow.

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Vehicle Budget

The table below lists cost by subsystem for one complete build out of the proposed design. The chart of subsystem budgets is comprehensive such that all subsystems necessary to mission completion are present with their respective budget totals. Itemized budgets of each subsystem appear in Appendix A.

Subsystem Cost Vehicle $765.22 Recovery $1383.35 Payload $1558.24 Propulsion $434.42 Total $4202.33 Table 60 Proposed Rocket Vehicle Budget Summary

Testing Budget

Testing is one of the most important areas in the development of any system, particularly a vehicle of this nature. This fact is not lost to the future engineers at Tarleton. The importance of testing a system is two-fold; it can eliminate product flaws and potential failures, and it can prove and improve upon a system’s overall safety. Taking the gravity of the task into consideration, the following budget was summarizes a preliminary testing budget. The following chart includes preliminary test budgets for each subsystem. Itemized test budgets are available in Appendix A.

Subsystem Unit Cost # of Test Unit Total Vehicle $765.22 3.5 $2,678.27 Recovery $1383.35 2 $2,766.7 Payload $1558.24 4 $6,232.96 Propulsion $434.42 5 $2,172.1 Total $13,850.03 Table 61 Preliminary Testing Budget Summary Outreach Budget

Outreach and community involvement events are critical to any engineering project both for publicity as well as educational opportunities. The project hosts many events that encourage students to participate in STEM fields. In order to host an effective outreach program a budget has been generated. The following chart includes a summarized budget of the outreach program.

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Event Cost Community Events $1,709.43 Scholastic Events $719.60 Equipment $140.74 Travel $1,100 Total $3,669.77 Table 62 Outreach Budget

Travel Budget

In order to meet all requirements of the SOW, the team must travel to Alabama and present their project. The logistics of moving a large number of personnel requires thought and careful planning in order to stay within budget. As of this report, the number of personnel traveling to Alabama is 15 and the required budget totals to $8200. The following chart consists of rough estimates of travel costs.

Service # of People Est. Cost Hotel 15 $4,200 Gas 15 $1,000 Van Rental 15 $1,000 Meals 15 $2,000 Total $8,200 Table 63 Estimated Travel Budget

Budget Summary

The following chart includes the projected budget for completing the project. The task of completing the NASA USLI is a complex interdisciplinary endeavor that tests the team’s knowledge and skills, including management of a budget. The first step in managing a budget is devising such a budget that is sufficient in meeting all costs necessary to complete the mission. Table 64 breaks down the known project costs, and a detailed budget is available for review in Appendix A.

Element Est. Cost Testing/Prototyping $13,972.23 Outreach $3,669.77 Final Build $4,202.33 Travel to Competition $8,200 Total $29,922.13

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Table 64 Preliminary Budget Summary

Funding Plan

A significant portion of the funding necessary for this project derives from a wide range of University organizations and other community support functions. Thus far, $11,500 in donations from the Tarleton President’s Circle, the Provost’s Office, the Dean of the College of Science, and the Tarleton Foundation fund the project. The Office of Student Research has provisions for the project amounting to $17,000. USLI Science Mission Directorate (SMD) funding also stems from NASA in the amount of $2,780. The total allocation for the project currently amounts to $31,280.

Figure 70 - Initial Funding Timeline

The Tarleton Aeronautical Team understands that a project of this magnitude requires a great deal of time and dedication. The following schedule to meet the requirements of the project serves as evidence in Figure 71. Gantt Charts detailing the project timeline follow. Gantt Timeline

The following chart gives a visual representation of major project deliverables.

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Figure 71 Project Timeline

Tarleton Aeronautical Team 145 2012 - 2013 USLI Preliminary Design Review

Testing Gantt Timeline

The Gantt chart below includes preliminary testing dates. Time is in the schedule to allow for lab prototyping and testing. The team plans to conduct multiple test launches including low altitude as well as high altitude test launches when possible. Using data gathered from these test launches, the team performs a failure analysis after each launch. These analyses are useful to optimize successive launches and overall design.

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Figure 72 Testing Timeline

Tarleton Aeronautical Team 147 2012 - 2013 USLI Preliminary Design Review

Outreach Gantt Timeline

The chart below delineates the dates the team plans for educational outreach. The star parties each involve a simple vehicle demonstration and a presentation about the basics of rocketry. During class trips, team members travel to area middle schools to actively engage students in safe, basic rocketry. The Tarleton Science Olympiad consists of area middle school and high school students convening at Tarleton to compete in science related activities. During the Science Olympiad the team demonstrates a static motor test in addition to giving presentations explaining basic rocketry. For more comprehensive information on the educational outreach component of the project, please refer to the following outreach section of this document.

Figure 73 Outreach Timeline

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Outreach Plan

Vehicle design, creation, and implementation are important components of this competition. Conjunctively, the educational engagement portion of this project is crucial, as its main goal is to promote enthusiasm for the necessary subjects that relate to rocketry and other important STEM fields. The team’s plan is to host several events for diverse audiences. All events aim to promote the global necessity of math, science, engineering, and technology. Furthermore, the team includes a vehicle launch with each event to provide a real world experience to reinforce the addressed STEM concepts in the lesson portion of each event. The team aims to encourage interest in the relevant subjects with the intent of increasing the number of people that choose to pursue STEM related careers.

Figure 74 - Acton Middle School

Educational Outreach

Educational outreach targets three main audiences through a variety of events; students, teachers, and the community as a whole. The team is currently establishing contacts and scheduling dates to visit the local middle schools. Several schools plan to participate in the educational outreach events already. Outreach to students in the schools occurs through a classroom lesson or an assembly style presentation. During the classroom sessions, small groups from the team present an original interdisciplinary lesson over rocketry with an emphasis on math and science.

The goal of these lessons is to demonstrate to students the importance of the STEM subjects and their role in a variety of topics such as engineering and rocketry. The necessity of these careers with companies such as NASA is a primary focus. By working in a setting which allows for a smaller student to presenter ratio, students are receive an opportunity to work closely with the team members on a lesson which reinforces concepts learned previously in a novel manner.

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Beyond the classroom lesson, the assembly format allows the team to communicate the same information to students on a larger scale. This portion of the outreach began at the request of some of the local middle schools. The assemblies take place toward the end of the school day and involve multiple classes and grade levels.

To conclude each school presentation, students join our team outside for a vehicle launch. The rocket launch adds to the lesson by giving the students a visualization of what they just learned. This increases students’ retention. In a classroom setting, interactive, hands-on lessons encourage learning.

Educator Outreach

By aligning the lessons created for the classroom presentations with state and national curriculum, these repeatable lessons remain relevant for reuse. The team is working to create a live webcast of a vehicle launch for teachers to access in their classes. This webcast allows teachers to use this online content as a real-life application in their classroom. Furthermore, the team is communicating with teachers throughout the state to distribute lesson plans. We hope to raise interest in STEM fields by having teachers join our group. Lesson plans are available from the Tarleton Aeronautical Team discussing rocketry and the importance of NASA.

Community Outreach

Star Party

Outreach beyond schools allows for students, teachers, and community members to join in learning about rocketry and STEM concepts. The team coordinates with Tarleton State University to co-host their Star Party event which occurs in both the fall and spring semesters. The event includes a discussion about the program, rocketry, the need for growth in the STEM fields, and a vehicle launch. The Star Party is an open invitation event; the team reaches audiences with a range from children to adults from local and surrounding areas.

Tarleton Regional Science Olympiad

The team will be at the eighth annual Tarleton Regional Science Olympiad on February 23, 2013. Students participating in this event along with their sponsors and family join the team for several vehicle launches including a static launch demonstration. A presentation and question and answer session follow. The day concludes with an awards ceremony. Again, the focus of this presentation is to promote the STEM fields and reiterate their importance pertaining to the nation’s progress.

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Participation Goal

The team expects to involve approximately 2,500 students in total. Comprehensive feedback from teachers and students will be gathered through surveys. This feedback helps the team alter presentations to ensure the quality of each event. Outreach efforts for boy scouts, girl scouts, and after school programs are being developed. The team members have a passion for the STEM fields, thus outreach is an important goal.

Accomplished Educational Outreach

On October 5, 2012 members of the Tarleton Aeronautical Team traveled to Granbury. Students at Acton Middle School and Granbury Middle School participated in basic rocketry presentations. The team’s presentation explained STEM fields and related careers. As part of Career Day at Acton Middle School, the team specifically discussed careers available at NASA. Bert led presentations featuring 7 Minutes of Terror, a NASA video highlighting interviews with NASA engineers on the Curiosity Rover project. While gaining exposure to career options with NASA, the students also learned the importance of safety protocol. To emphasize the message of the video, the students were given the opportunity to experience rocketry in a safe environment. Before the team launched their rocket, each class was given the opportunity to launch two-liter water bottle rockets. The surveys reflected that the majority of the students were delighted with their experience. In the survey, the students were asked to report whether they felt a greater interest in Science, Mathematics, Engineering, or Technology, three things they learned from the presentation, and what their favorite parts of the presentation were. The data is given in Table 65 and illustrated in Figure 75.

Subject Interest Count Science 39 Math 35 Engineering 36 Technology 33 Rocketry 14 None 15 Table 65 Accomplished Educational Outreach

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Subject Interest

Science Math Engineering Technology Rocketry None 9% 8% 23%

19%

20% 21%

Figure 75 Subject Interest

Areas of learning include force concepts, propulsion, rocket construction, rocket design, launch procedures, failure modes, qualifications for building rockets, careers in rocketry, competitions in amateur rocketry, and information about NASA. The categories with the greatest percentage of student learning were rocket design, qualifications for building rockets, launch procedures, and propulsion. This indicates that more time should be spent in future presentations on the other learning categories, but further sampling is required. The data is given in Table 66 and illustrated in Figure 76.

Presentation Learning Outcomes Count Forces 5 Propulsion 22 Construction 8 Design 31 Procedures 21 Failure Modes 13 Qualifications 24 Careers 14 Competitions 5 NASA 11 None 8 Table 66 Presentation Learning Outcomes

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Presentation Learning Outcomes

Forces Propulsion Construction Design Procedures Failure Modes Qualifications Careers Competitions NASA None 5% 3% 7% 13% 3% 9% 5%

19% 15%

8% 13%

Figure 76 Presentation Learning Outcomes When asked what their favorite part of the presentation was, the greatest number of students responded in favor of the water bottle rocket activity. The data is given in Table 67 and illustrated in Figure 77.

Favorite Part Count Launch 11 Launch Failure 3 Flight 5 Outside 10 Water Rockets 27 Big Rocket 21 None 7 Table 67 Favorite Part

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Favorite Part

Launch Launch Failure Flight Outside Water Rockets Big Rocket None 8% 13% 4% 25% 6%

12%

32%

Figure 77 Favorite Part

The surveys conducted at the two middle schools on October 5, 2012 were free response. The Granbury events were the first conducted by the team. They provided a wide variety of student responses concerning the presentation and demonstration. This feedback will ultimately be used to formulate a comprehensive and unbiased multiple-choice survey to be conducted at subsequent events. This will boost the quality of questions posed at future presentations.

Conclusion

The team is eager to continue the design process and see the final product come to fruition. Up to this point, the project is both challenging and rewarding. Although this is the first time to compete in NASA’s USLI, the Tarleton Aeronautical Team is confident in its ability to design a vehicle and payload to the customer prescribed specifications. Creative, legitimate solutions have been posed and implemented, defining the performance characteristics of the vehicle and payload systems as a whole.

The vehicle design features reflect standard practices in amateur rocketry. All components of the vehicle aim to achieve the goal of delivering the SMD payload to one mile above ground level.

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The SMD payload criterion calls for an atmospheric data gathering instrument that meets all requirements as stated in the SOW. Controlling the sensors to make measurements is among a long list of complicated problems creating a fully functional payload. Other taxing problems include the orientation of the image camera and the placement of the UV and solar irradiance sensors. The self- leveling camera design aims to take proper images by correcting offsets in the payload orientation during descent. Pyranometers and UV sensors sit in an opposing manner to effectively increase the field of view for solar irradiance and ultraviolet measurement. Optical properties of the clear acrylic payload housing must be addressed, and sensor performance testing must occur to verify this housing structure selection. The payload will ultimately be inserted in a PCB due to advantages in size, efficiency, repeatability and overall fidelity in design.

The educational outreach portion of the project is going extremely well. The minimum requirement for the number of students to be reached was exceeded on first day of this competition. The team continues to go above and beyond this minimum requirement. The intention of the team is to expose as many students to the STEM fields as possible.

Building upon the momentum spawned by the success in the 2012 CanSat competition, along with support from the community and University, the team is eager to progress through the design life cycle. Ultimately, flying the final vehicle and payload on launch day will illustrate the team’s achievements. It is a true testament to the abilities and ingenuity of the team members, and provides an invaluable exercise in creating real-world engineering experience.

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Appendix A - Itemized Subsystem Budget A.1 Structure System

Body Part Number Seller Price Per # of Units Total Unit

Nose Cone FNC5.5EL WildMan $47.50 1 $47.50 Rocketry

Fiberglass Body G12-5.5-60 WildMan $38.84 6 $233.04 Tube Rocketry

Acrylic Body Tube ACRCAT5.500ODX.250 ePlastics $39.05 2 $78.10

Sheet for Fins 500SHT0.125X48X96 ePlastics $208.00 0.2 $41.60

Motor Tube G12-3.0-48 WildMan $71.06 1 $71.06 Rocketry

Bulk Plate PVCGRAY2.00LAM12x24 ePlastics $147.82 0.25 $36.96 (Payload)

Bulk Plate Standard FBP5.5 WildMan $7.60 5 $38.00 Rocketry

Couplers G12CT-5.5 WildMan $47.03 2 $94.06 Rocketry

Centering Rings FCR5.5-3.0 WildMan $8.55 3 $25.65 Rocketry

Epoxy 4500Q WildMan $69.00 0.25 $17.25 Rocketry

Motor Retainer RA75 Wildman $52.00 1 $52.00 Rocketry

Shipping $15.00 2 $30.00

Grand $765.22 Total

A.2 Recovery System Budget

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Proposed Selection Distributor Item Number Unit Proposed Typical Cost Quantity Unit Cost

Main Altimeters Featherweight Raven3 $155.00 2 $155.00 Altimeters

Backup Altimeters PerfectFlite StratoLogger $79.95 3 $79.95

Electric Matches Coast Rocketry Daveyfire N28BR $2.95 4 $11.80

FFFFg Black Powder Goex Goex 4F Black $15.75 1 $15.75 Powder

Black Powder Ejection Aerocon Systems BPSmall $3.00 1 $3.00 Charge Holders

Swivels Commonwealth SWLDK80 $1.99 2 $3.98 Rocketry

Main Shock Cord Giant Leap Tubular Kevlar $37.99 1 $37.99 Rocketry

Drogue Shock Cord Giant Leap Tubular Kevlar $31.49 1 $31.49 Rocketry

Main Parachute Recovery Premium with Spill $145.00 1 $145.00 Technologies Hole

Flameproof Main Rocketman DB8 $40.00 1 $40.00 Parachute Enterprises Deployment Bag

Drogue Parachute Recovery Premium with $16.00 1 $16.00 Technologies Spillhole

Flame-Proof Drogue Rocketman DB2 $25.00 1 $25.00 Parachute Enterprises Deployment Bag

U-Bolts Sunward U-Bolt Assembly - $4.29 2 $8.58 Aerospace 0.25in. (compact)

Quick Links Commonwealth 0.25in. Stainless $2.99 4 $11.96 Rocketry Steel Delta Quick Link

Shear Pins Missile Works 2-56 Nylon Shear-Pin $1.00 1 $1.00 (10 pack)

Arming Switches Featherweight Featherweight $25.00 5 $125.00 Altimeters Magnetic Switch

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Hand-held Rotary Fan WeatherShack SpeedTech WM-300 $154.95 1 $154.95 Anemometer

GPS Big Red Bee BeeLine GPS- $289.00 2 $578.00 Package Deal

Total $1,444.45

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A.3 Payload Budget

Product Price Quantity Total

Arduino 2560-R3 $58.95 1 $58.95 micro SDHC $9.99 2 $19.98

Adafruit 254 $15.00 2 $30.00

BMP180 $15.00 1 $15.00

MS5611-01BA03 $29.99 1 $29.99

HIH4030 $16.95 1 $16.95

HH10D $9.95 1 $9.95

SP-110 $169.00 2 $338.00

TSL2561 $12.50 2 $25.00

SU-100 $159.00 2 $318.00

TOCON_ABC3 $148.00 2 $296.00

HS-85BB+ Mighty Micro $19.99 2 $39.98

Adafruit 397 $42.00 1 $42.00

VCC-003-MUVI-BLK $49.95 1 $49.95

Xbee-PRO XSC S3B $42.00 1 $42.00

LS20031 $60.00 1 $60.00

LCD-11062 $34.95 1 $34.95

ADXL345 $27.95 1 $27.95

Ultralife U9VLBP $6.65 8 $53.20

Mouser 534-3427 $10.22 2 $20.44

Adept A1E $29.95 1 $29.95

$1,558.24

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