flight readiness review - tarleton.eduflight readiness review i note to reader: to facilitate the...

[Type text] I) Summary of FFR Report [Type text]

Flight Readiness Review

2012-2013 NASA USLI

“We can lick gravity, but sometimes the paperwork is overwhelming.” - Werner Von Braun

Tarleton State University Flight Readiness Review

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Note to reader:

To facilitate the reading of the Flight Readiness Review, we have mirrored the Student

Launch Project Statement of Work. In the body of the FRR, you will find extensive detail

in the design of our SMD payload. The payload‟s features are threefold; atmospheric

data gathering sensors, a self-leveling camera system, and a 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 250 man hours of work. The PCBs provide major enhancement of the

signal integrity of the sensor data in addition to their space and power efficient qualities.

We have enjoyed finalizing our design for flight readiness and submit this document for

your review.


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Table of Contents Flight Readiness Review ...................................................................................................

2012-2013 NASA USLI ..................................................................................................

I) Summary of FFR Report .............................................................................................. 1

Team Summary ........................................................................................................... 1

Launch Vehicle Summary ............................................................................................ 1

Payload Summary ........................................................................................................ 1

II) Changes Made since CDR .......................................................................................... 2

III) Vehicle Criteria ........................................................................................................... 4

Design and Construction of the Vehicle ....................................................................... 4

Structural Elements .................................................................................................. 4

Drawings and schematics ....................................................................................... 13

Flight Reliability Confidence and Mission Success Criteria .................................... 16

Testing .................................................................................................................... 17

Workmanship .......................................................................................................... 18

Vehicle Safety and Failure Analysis ........................................................................ 19

Full-scale Launch Results ....................................................................................... 19

Mass Report ........................................................................................................... 27

Recovery Subsystem ................................................................................................. 28

Design Defense ...................................................................................................... 28

Parachute Justification ............................................................................................ 43

Safety and Failure Analysis .................................................................................... 43

Mission Performance Predictions ............................................................................... 50

Mission Performance Criterion................................................................................ 50

Kinetic Energy Management ................................................................................... 59

Altitude and Drift Predictions .................................................................................. 60

Verification of System Level Functional Requirements ........................................... 60

Safety and Environment ............................................................................................. 65

Safety and Mission Assurance Analysis ................................................................. 65

Payload Integration ................................................................................................. 78

Integration of the Payload with the Launch Vehicle ................................................ 78

Compatibility of Elements ....................................................................................... 81

Payload Interface Dimensions ................................................................................ 82


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Payload-Housing Integrity ....................................................................................... 82

Integration Demonstrated ....................................................................................... 83

IV) Payload Criteria ....................................................................................................... 86

Experiment Concept .................................................................................................. 86

Science Value ............................................................................................................ 86

Payload Objectives ................................................................................................. 86

Payload Success Criteria ........................................................................................ 86

Test and Measurement, Variables, and Controls .................................................... 89

Payload Design .......................................................................................................... 90

ADGS ..................................................................................................................... 93

ARTCOS ................................................................................................................. 99

Precision of Instrumentation and Repeatability of Measurements ........................ 104

Flight Performance Predictions ............................................................................. 107

Test and Verification Program .............................................................................. 116

Payload Requirement Verification ............................................................................ 117

Verification Statements ......................................................................................... 118

Safety and Environment (Payload) .......................................................................... 119

Personnel Hazards ............................................................................................... 122

Environmental Concerns ...................................................................................... 123

V) Launch Operations Procedures .............................................................................. 124

Checklists................................................................................................................. 124

Recovery Preparation Checklist (to be completed before arriving to launch site): 124

Motor Preparation Checklist: ................................................................................ 125

Igniter Installation Checklist: ................................................................................. 126

Launchpad Setup Checklist: ................................................................................. 126

Launch Procedures Checklist: .............................................................................. 127

Troubleshooting Checklist: ................................................................................... 128

Post-flight Inspection Checklist: ............................................................................ 129

Payload Preparation (to be completed before arriving at launch field): ................. 129

Pre-launch Payload Preparation and Ground Station Setup: ................................ 130

Safety Materials Checklist .................................................................................... 130

Safety Checklist .................................................................................................... 130

The following is an itemized components checklist for each subsystem. ................. 131

Structure Components: ......................................................................................... 131


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Recovery Components: ........................................................................................ 131

Payload and Ground Station Components: .......................................................... 132

Safety and Quality Assurance .................................................................................. 132

Risk Assurance ..................................................................................................... 132

Risk Assessment for Launch Operations .............................................................. 132

Environmental Concerns ...................................................................................... 134

Identification of person responsible ...................................................................... 134

VI) Project Plan ........................................................................................................... 135

Budget Summary ..................................................................................................... 135

Funding Plan......................................................................................................... 142

Timeline ................................................................................................................... 142

Outreach Timeline ................................................................................................ 146

Education plan ......................................................................................................... 146

Outreach Plan ....................................................................................................... 146

Accomplished Educational Outreach .................................................................... 149

VII) Conclusion ............................................................................................................ 161


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Table of Figures Figure 1: Launch Vehicle Dimensions ............................................................................. 4 Figure 2: Nose Cone ....................................................................................................... 5 Figure 3: Ballast and GPS mount .................................................................................... 6 Figure 4: Coupler Frame ................................................................................................. 6 Figure 5: Coupler ............................................................................................................ 7 Figure 6: Avionics Bay ..................................................................................................... 7 Figure 7: Avionics Bay Lid ............................................................................................... 8 Figure 8: Payload Bulkhead and Avionics Rail System ................................................... 8 Figure 9: Avionics Bay Outside of Coupler ...................................................................... 9

Figure 10: Avionics Sled Layout ...................................................................................... 9 Figure 11: Acrylic Payload Housing Structure ............................................................... 10

Figure 12: Coupler Screw in Attachment ....................................................................... 10 Figure 13: Weight Simulator .......................................................................................... 11 Figure 14: Booster Section ............................................................................................ 11 Figure 15: Fin Dimensions ............................................................................................ 12

Figure 16: Shock Cord Epoxied to Motor Mount Tube .................................................. 12 Figure 17: Dual Attachment Scheme............................................................................. 13 Figure 18: Booster Section and Motor Mount Tube ....................................................... 14

Figure 19: Upper Body Airframe with Nose Cone ......................................................... 14 Figure 20: Payload Housing Structure, Aluminum Framework, and Couplers ............... 15

Figure 21: Vehicle Assembly ......................................................................................... 15 Figure 22: Rendereing of Final Vehicle ......................................................................... 16 Figure 23: Simulation Data ............................................................................................ 20

Figure 24: OpenRocket Rendering of Vehicle ............................................................... 20

Figure 25: Simulated Flight Data ................................................................................... 20 Figure 26: Drogue Stratologger Flight Data ................................................................... 22 Figure 27: Main Stratologger Flight Data ....................................................................... 22

Figure 28: Simulation Data ............................................................................................ 23 Figure 29: OpenRocket Rendering of Vehicle ............................................................... 23

Figure 30: Simulated Flight Data ................................................................................... 24 Figure 31: Drogue Stratologger Data ............................................................................ 25 Figure 32: Drogue Raven3 Flight Data .......................................................................... 25 Figure 33: Main Stratologger Flight Data ....................................................................... 26

Figure 34: Main Raven3 Flight Data .............................................................................. 26 Figure 35: Packed Deployment Bag .............................................................................. 33 Figure 36: Deployment Bag Operation .......................................................................... 33

Figure 37: Altimeter Wiring Configuration ...................................................................... 37 Figure 38: Stratologger Software Flow Diagram ........................................................... 38 Figure 39: GPS Software Flow Diagram ....................................................................... 39 Figure 40: Avionics Bay Structural Components ........................................................... 40

Figure 41: Attachment Bulkhead Inside Avionics Coupler ............................................. 40 Figure 42: Couplers in Profile ........................................................................................ 41 Figure 43: Structural Assembly of the Avionics Bay ...................................................... 41 Figure 44: GPS Device and Wireless Remote .............................................................. 42 Figure 45: Final Vehicle Simulation ............................................................................... 52


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Figure 46: Input Parameters for Final Simulation .......................................................... 52

Figure 47: L1720-WT Actual Thrust Curve .................................................................... 53 Figure 48 - PDF of Apogee for Competition Flight ........................................................ 55 Figure 49 - PDF of Lateral Distance for Competition Flight ........................................... 55 Figure 50: Center of Gravity and Center of Pressure .................................................... 57 Figure 51: Rear Payload Bulkhead ............................................................................... 79

Figure 52: Bracket on Forward Payload Bulkhead ........................................................ 79 Figure 53: Securing Payload Framework to Rear Bulkhead .......................................... 80 Figure 54: Framework Fitting into the Bracket ............................................................... 80 Figure 55: Fully Assembled SMD Payload Section ....................................................... 81 Figure 56: Payload Framework Secured to Rear Payload Bulkhead ............................ 83

Figure 57: Payload Framework Secured ....................................................................... 84 Figure 58: Payload Integration ...................................................................................... 84

Figure 59: Payload Integration ...................................................................................... 84

Figure 60: Integrated Payload ....................................................................................... 85 Figure 61: Payload Design ............................................................................................ 91 Figure 62: Constructed Payload .................................................................................... 92 Figure 63: ADGS PCB Schematic ................................................................................. 94

Figure 64: ADGS PCB ................................................................................................... 95 Figure 65: Power PCB Schematic ................................................................................. 96

Figure 66: Power Board ................................................................................................ 97 Figure 67: Solar Sensor Array ....................................................................................... 98 Figure 68: Real Time Sensor Readings ........................................................................ 98

Figure 69: ARTCOS Control PCB Schematic ................................................................ 99 Figure 70: ARTCOS Powerboard PCB Schematic ...................................................... 100

Figure 71: ARTCOS Explosion Assembly ................................................................... 101 Figure 72: ARTCOS Assembly .................................................................................... 102

Figure 73: Component Connections ............................................................................ 103 Figure 74: SU-100 Spectral Response ........................................................................ 105

Figure 75: Acrylic Sheet UV Transmission Curves ...................................................... 105 Figure 76: Spectral Response for Apogee Pyranometer ............................................. 106 Figure 77: Pressure Data ............................................................................................ 107

Figure 78: CDR Pressure Data ................................................................................... 108 Figure 79: Launch Pad Temperature........................................................................... 108 Figure 80: Temperature Data ...................................................................................... 109

Figure 81: CDR Launch Pad Humidity ........................................................................ 110 Figure 82: Lux Sensor Data ........................................................................................ 110

Figure 83: Solar Irradiance .......................................................................................... 111 Figure 84: CDR Solar Irradiance Data......................................................................... 112 Figure 85: GPS Flight Data ......................................................................................... 113 Figure 86: CDR GPS Flight Data ................................................................................ 114 Figure 87: ARTCOS Flight and Landing Photos .......................................................... 115

Figure 88: President's Circle Awarding the Aeronautical Team ................................... 142 Figure 89: Gantt Testing Timeline ............................................................................... 145 Figure 90: Outreach Timeline ...................................................................................... 146

Figure 91: Students Launching a Water Rocket .......................................................... 147


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Figure 92: Team Members Educate and Entertain Acton Students ............................ 149

Figure 93: Educational Outreach Survey Results ........................................................ 150 Figure 94: Presentation Learning Outcomes ............................................................... 151 Figure 95: Students' Favorite Portion of Presentations ............................................... 152 Figure 96: Students Received Stickers for Correct Responses ................................... 155 Figure 97: Students Engaged in Learning ................................................................... 156

Figure 98: Team Members Addressed Large Audiences of Students ......................... 157 Figure 99: Students Learning at the Recovery Station at Dublin Middle School ......... 159 Figure 100: Students Enjoying the Art Station, Decorating Parachutes ...................... 159


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Index of Tables Table 1: Vehicle Size and Mass ...................................................................................... 1 Table 2: Experiment Summary ........................................................................................ 1 Table 3: Changes Made to Vehicle Criteria ..................................................................... 2 Table 4: Changes Made to Recovery .............................................................................. 3 Table 5: Launch Conditions ........................................................................................... 19 Table 6: Simulated Flight Data ...................................................................................... 21 Table 7: Recovery Electronics Data .............................................................................. 21 Table 8: Launch Conditions ........................................................................................... 23

Table 9: Simulated Flight Data ...................................................................................... 24 Table 10: Recovery Electronics Data ............................................................................ 24

Table 11: Mass Summary ............................................................................................. 27 Table 12: Mass by Subsection ...................................................................................... 28 Table 13: Test Flight 1/26/13 ......................................................................................... 44 Table 14: Test Flight 2/14/13 ......................................................................................... 45

Table 15: Test Flight 2/19/13 ......................................................................................... 46 Table 16: Test Flight 3/7/13 ........................................................................................... 47 Table 17: Test Flight 3/12/13 ......................................................................................... 48

Table 18: Test Flight 3/14/13 ......................................................................................... 49 Table 19: Test Flight 3/15/13 ......................................................................................... 50

Table 20 - Confidence Intervals for Apogee and Drift.................................................... 54 Table 21: Calculated versus Simulated CG and CP Measurements ............................. 59 Table 22: Kinetic Energy by Section During Flight ........................................................ 60

Table 23: Altitude Predictions ........................................................................................ 60

Table 24: Drift Predictions ............................................................................................. 60 Table 25: Vehicle Verification Table .............................................................................. 64 Table 26: Potential Failure Modes for the Design of the Vehicle ................................... 67

Table 27: Potential Structure Failure Modes and RPNs ................................................ 68 Table 28: Potential Propulsion Failure Modes aand RPNs ............................................ 69

Table 29: Potential Recovery Failure Modes and RPNs ............................................... 73 Table 30: Top 10 Failure Modes by RPN ...................................................................... 74 Table 31: Potential Hazards to Personnel ..................................................................... 75 Table 32: Legal Risks .................................................................................................... 77

Table 33: Effects of Materials used in Construction and Launch ................................... 78 Table 34: Environmental Factors Hendering Recovery ................................................. 78 Table 35: Procedures for Installing Payload .................................................................. 81

Table 36: Payload Interface Dimensions ....................................................................... 82 Table 37: Payload Objectives ........................................................................................ 87 Table 38: FR-4 Characteristics ...................................................................................... 93 Table 39: Payload Electrical Components ................................................................... 104

Table 40: Payload Sensor Precision ........................................................................... 106 Table 41: Payload Functional Requirements ............................................................... 118 Table 42: Potential Failure Modes for Payload Section During Integration ................. 119 Table 43: Potential Failure Modes for Payload During Launch Operations ................. 121 Table 44: Top Five Potential Failure Modes by RPN .................................................. 121


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Table 45: Potential Hazards to Personnel ................................................................... 123

Table 46: Risk Assessment for Launch Operations .................................................... 134 Table 47: Known Project Costs ................................................................................... 135 Table 48: Team Budget ............................................................................................... 135 Table 49: Structure/Propulsion System Budget ........................................................... 136 Table 50: Recovery System Budget ............................................................................ 137

Table 51: Payload Budget (Through-Hole PCB) ......................................................... 138 Table 52: Payload Budget (PCB/Non PCB) ................................................................ 141 Table 53: Educational Outreach Survey Results ......................................................... 150 Table 54: Presentation Learning Outcomes ................................................................ 151 Table 55: Students' Favorite Portion of Presentations ................................................ 152

Table 56: Educational Outreach Stations .................................................................... 155 Table 57: Educational Outreach Stations .................................................................... 158

I) Summary of CDR Report

I) Summary of FFR Report

Team Summary Tarleton Aeronautical Team Tarleton State University Box T-0470 Stephenville, Texas 76402 Team Mentor: Pat Gordzelik. Past and Present Credentials: Tripoli Amarillo #92 Board of Directors Member, Technical Advisor Panel Panhandle of Texas Rocketry Society Inc. – Founder, President, Prefect TRA 5746 L3 NAR 70807 L3CC Committee Chair Married to Lauretta Gordzelik, TRA 7217, L2.

Launch Vehicle Summary Size and Mass

Length 97.5 inches

Outer Diameter 5.525 inches

Mass 35.9 pounds

Motor

Selection Cesaroni L1720-WT-P

Recovery

Drogue 24” Silicone Coated Rip stop Nylon Parachute, Apogee Deployment

Main 120” Silicone Coated Rip stop Nylon Parachute, 700 foot AGL Deployment

Avionics Primary PerfectFlite Stratologger Altimeter, Backup PerfectFlite Stratologger Altimeter, and Garmin GPS Tracking

Rail Size

Rail 1010 and 1515

Milestone Review Flysheet – see link on website

Table 1: Vehicle Size and Mass

Payload Summary

Title Experiment

Science Mission Directorate (SMD) Payload

Sponsored by NASA; Gather Atmospheric and GPS Data, Autonomously Orient Photographic Camera, Capture Video for Public Outreach

Table 2: Experiment Summary


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II) Changes Made since PDR

II) Changes Made since CDR

Changes Made to Vehicle Criteria Structure Rationale

Avionics Port Holes at 5/32 inch Stratologger Manual Specification

Booster Length is 32 inches (approved*) Weight Reduction

GPS location to nosecone Poor attachment scheme on shock chord

Payload Housing is 27 inches (approved*) Weight reduction

Avionics Sled Guides Changed Aluminum material strength

Payload Framework is 20.5 inches Payload integration

Payload Framework to Bulkhead Mount (Bracket) Payload integration

Payload Framework is 0.5 by .75 inch aluminum angle Added strength

*These changes were submitted via email directly after CDR on 3/4/2013

Recovery Rationale

4 Strattologgers are used for all Altimeters Power problems with Raven 3s

1 battery is used for each avionics bay Power indicator

1 LED per altimeter Power Indicator

1 second delay set on backup apogee altimeter Redundancy

700 foot main parachute deployment setting; 650 foot backup setting

Redundancy

J-Tek 10 E-matches with plastic test tubes Reliability

Ejection charges use clip connectors Safety

Kevlar TAC-9B Deployment Bag Ease of Packing

24 inch Pilot drogue for deployment bag Required for Deployment bag

Deployment Bag Attachment Scheme Required for Deployment bag

Recovery Rationale 1/4 inch to 3/8 inch tubular Kevlar shock chord (approved**)

Greater FOS; Packing constraints

Main Parachute packing scheme Prevent entanglement

**RSO approved on 3/13/2013

Changes Made to Payload Criteria

Payload Rationale

LCD Screen removed Problems in testing, power consumption

4th battery added Power Activation Circuit

Table 3: Changes Made to Vehicle Criteria


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II) Changes Made since PDR

4 UV sensors 90 degree intervals; better data acquisition

4 Solar Irradiance 90 degree intervals; better data acquisition

Minor PCB Changes Layout Optimization

The team made no significant changes to the project plan.

Table 4: Changes Made to Recovery


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III) Vehicle Criteria


Design and Construction of the Vehicle The entire launch vehicle, as shown in Figure 1, is 97.5 inches long. The fin span measures 13.525 inches. This includes the 5.525 inch width of the airframe. Each section of the launch vehicle will be described in more detail in the following subsections.

Structural Elements

Upper Body Airframe This section contains the nose cone and upper body airframe.

Airframe

The upper body airframe is 28.0 inches long. It has a 5.525 inch outside diameter and 5.375 inch inside diameter. These dimensions were chosen in order to fully house the payload and main recovery system. Fiberglass was chosen as the material due to its

Figure 1: Launch Vehicle Dimensions


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strength and durability. This allows it to withstand the forces experienced during acceleration and landing.

Nose Cone

The elliptical nose cone (Figure 2) is attached to the upper body airframe with four screws and drywall anchors. It is 8.5 inches long and has a 4.6 inch shoulder. Fiberglass was chosen for the nose cone for its strength and durability.

Ballast system/Attachment/GPS

The ballast system is built into the upper body airframe and contained in the shared space of the upper body airframe and the nose cone. A bulkhead is epoxied into the airframe where the shoulder of the nose cone sits. A hole is drilled through the center of the bulkhead and a welded 5.5 inch eyebolt is secured through it. The eyebolt serves as the mounting point for the main parachute on the opposite side of the bulkhead. It is important to note that all eyebolts used for attachment hardware are welded to increase the maximum load rating. The threaded portion of the eyebolt serves as a location to secure both the ballast weights and the Garmin Astro DC40 GPS. An assortment of washers serves as the ballast weight as seen below in Figure 3. To secure different sized washers, they are staggered in different directions and use a spacer to distribute the force.

Figure 2: Nose Cone


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Couplers The couplers are handmade to fit the SMD payload housing to both the upper body airframe and the booster section. The coupler frame can be seen in Figure 4. Each section will remain attached to the payload housing throughout the entire flight. Commercially available couplers are not used because of the varying inside diameters of the fiberglass and the acrylic sections. Thus, two different couplers are overlapped to make one single coupler measuring 11.25 inches in length. One side of the coupler is 5.372 inches in diameter and the other side is 5.248 inches in diameter, which can be seen in Figure 4.

Each coupler has a collar of fiberglass epoxied into place where the two sections meet. In Figure 5, the fiberglass collar goes where the purple section meets the red section. Four 0.15625 inch portholes are located on the collar to allow for easier vehicle assembly by preventing any misalignment. An LED is also mounted to the coupler on the collar. The LED provides visual verification when the avionics switch is activated.

Figure 3: Ballast and GPS mount

Figure 4: Coupler Frame


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Shear pins are used in the coupler between the upper body airframe and SMD payload housing to prevent premature separation at apogee when the drogue parachute‟s charges fire.

Avionics Bays

The avionics bays are housed within each coupler. They are separated from the SMD payload by bulkheads and the other sections by bay lids. Each avionics sled is supported by a rail system consisting of two 10 inch rods of 0.25 inch steel all thread. The inside of each avionics bay is lined with aluminum flashing to shield the avionics from outside interference, as seen in Figure 6.

Each avionics bay lid, shown in Figure 7, consists of two fiberglass bulkheads. One bulkhead fits the inside diameter of the airframe, and the other fits the coupler to keep the bulkhead centered. This prevents the lid from shifting and becoming lodged within the bay. A welded eyebolt is mounted in the center and used as an attachment point for the recovery system. Two 0.25 inch holes are drilled through the lid on either side of the eyebolt to slide onto the rail system. A 0.125 inch hole is drilled through the lid to allow the ejection charge leads from the altimeters to pass.

Figure 5: Coupler

Figure 6: Avionics Bay


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The bulkheads separating each avionics bay from the SMD payload housing were machined out of PVC using a lathe. They are each one inch thick with material lathed out to leave a cup shape on the side facing the avionics. Six 0.1875 inch holes are drilled through the side of the bulkhead at 60 degree intervals for screws. Two 0.125 inch holes are offset from each screw hole to secure self locking nut plates inside the bulkhead using rivets. Two 0.25 inch holes are drilled through the bulkhead for the rail system, which is secured using washers, lockwashers, and nuts. This design is displayed in Figure 8.

Figure 7: Avionics Bay Lid

Figure 8: Payload Bulkhead and Avionics Rail System


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Each avionics sled is made out of 0.25 inch thick plywood. Aluminum tubes are attached to the sled using coax cable mounts as shown in Figure 9. These are used to easily slide the sled onto the rail system of the avionics bay. The layout of the avionics sleds is displayed in Figure 10. Each component is mounted to the sled using screws and nuts. The battery is further secured to the battery mount using cable ties to prevent them from becoming dislodged during flight.

SMD Payload Housing The payload housing structure is 27 inches long. It is made of clear acrylic for visible verification that the payload is powered on and functioning as illustrated in Figure 11. This also allows UV sensors and cameras to be mounted internally thus not requiring payload ejection. As seen in Figure 12, the couplers are mounted with six screws containing nylon grommets to prevent the metal screws from contacting the acrylic. The

Figure 9: Avionics Bay Outside of Coupler

Figure 10: Avionics Sled Layout


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payload mounting rail is attached to one coupler with screws and held in place by a bracket on the other. Portholes for the payload are located on the acrylic eight inches from upper body airframe. Five 0.25 inch holes are drilled to provide proper venting.

Mass Simulator

For the full-scale flight, the payload was omitted and a mass simulator was used in place. A similar payload mounting rail was lined with aluminum framework and washers were attached to simulate the weight of the components of the payload. Weights were placed as shown to accurately represent mass locations of the actual payload. The simulator weighed 2.59 pounds. The completed simulator can be seen in Figure 13.

Figure 11: Acrylic Payload Housing Structure

Figure 12: Coupler Screw in Attachment


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Booster section The booster section of the vehicle is 28 inches long. Mounted inside the booster section is a 20 inch motor mount tube as pictured in Figure 14. The motor mount tube is three inches in diameter to accommodate a 75 millimeter motor. Slots measuring 0.125 inch wide are cut into the booster section starting 1.125 inches from the bottom of the airframe and extending 9.7 inches for the fin tabs.

Each fin tab is made of 0.125 inch thick fiberglass and is secured with epoxy. The fiberglass can has high durability and increases the chance that they can withstand impact and be reusable. The fin dimensions can be seen in Figure 15.

Figure 13: Weight Simulator

Figure 14: Booster Section


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For safety of the recovery system, the booster section contains dual recovery attachment. The drogue parachute attaches via shock cord to a larger 0.375 inch welded eyebolt attached to the motor casing as well as a shock cord epoxied along the entire length of the motor mount tube as seen in Figure 16. The dual attachment scheme is shown in Figure 17.

Figure 15: Fin Dimensions

Figure 16: Shock Cord Epoxied to Motor Mount Tube


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Drawings and schematics

The vehicle design is illustrated in the following figures. These show the various internal characteristics to each section in light gray lines. Figure 18 shows the booster section and motor tube layout.

Figure 17: Dual Attachment Scheme


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Figure 19 shows the layout of the upper body airframe with nose cone attached.

Figure 20 shows the payload housing structure with aluminum framework installed and couplers that house the avionics bays.

Figure 18: Booster Section and Motor Mount Tube

Figure 19: Upper Body Airframe with Nose Cone


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Figure 21 shows the layout of the entire vehicle assembly.

Figure 22 is a CAD rendering of the final vehicle.

Figure 20: Payload Housing Structure, Aluminum Framework, and Couplers

Figure 21: Vehicle Assembly


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Flight Reliability Confidence and Mission Success Criteria The confidence in the design of the vehicle is based off of two major factors, testing and assembly. Careful assembly and precision in manufacturing provide confidence that vehicle will withstand all loading during launch and throughout the flight. With the high number of test launches conducted, the structural integrity and aerodynamic profile of the vehicle has proven very reliable. Considering motor choice, fin and nose cone shape, the resulting stability margin, total weight, and the final design of the launch vehicle has demonstrated that a safe and reliable flight can occur. The project defines the mission as a vehicle flight with a payload onboard where both 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 of altitude, and the official scoring altimeter will be intact, audible, and report altitude. The recovery system stages a deployment of the drogue parachute at apogee and deploys the main parachute at 700 feet. After apogee and descent, the entire vehicle lands within 2,500 feet of the launch pad. The vehicle design is fully verified in Table 25 If the above conditions are met, the mission will be considered partially successful in that requirements have been met by the vehicle design. However, because the actual altitude of the vehicle at apogee is scored based on comparison to one mile above ground level, a successful mission would be warranted only if the aforementioned conditions are met and an apogee of exactly 5,280 feet is achieved, plus or minus 0.1% plus one foot due to precision of the scoring altimeter.

Figure 22: Rendereing of Final Vehicle


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Testing

Static testing on individual components, such as airframe strength, was completed prior to CDR. The results of the static tests were conclusive that each component was strong enough to withstand all expected loads, both during liftoff and at landing. This has been further verified through the large number of test flights conducted to date. Since CDR, 7 test flights have demonstrated a very durable vehicle design. Functional testing has been achieved through extensive test flights. Sub-scale flight testing was used to minimize losses while testing the parachute attachment scheme and avionics. Once each subsystem was perfected, the first prototype was built to perfect the integration of all components and test the design of the airframe and aerodynamics. Before the first prototype was completed, simulation using OpenRocket provided a method of testing the size, shape, and weight of the rocket. Each component in the final build must work together perfectly for a successful flight and recovery. This was demonstrated in the first full-scale launch of the final build. This launch contained a mass simulator for the payload. The flight was within 30 feet of the simulation, and was recovered with no damages. This flight provided verification of a successful design, assembly, and launch procedure. As documented in the CDR, the results of the static tests on the vehicle components are included below. All Structure Tests were done using a hydraulic press. Force is applied in ~10 pound increments. At every 100 pound increment, the press was released, and then reapplied to that weight instantly to represent shock force. The scale used to measure the force was an airplane scale with a maximum of 1,500 pounds. These tests were done at the Polen Facility in Granbury, Texas. The tests were done in a way to minimize destruction to both the components and the facility. They were overseen by facility owner Richard Keyt, a former Air Force pilot and licensed aircraft mechanic/machinist who holds a Bachelor of Science degree in Aeronautical Engineering from the University of Minnesota.

Airframe To test the strength of the 0.125 inch thick bulkhead, a 2 inch by 3 inch block was placed on a bulkhead and a hydraulic press was used to apply force to it. The result was that it held the maximum amount of weight, 1500 pounds, that the scale could measure. Using this number and dividing by the contact area, the flat sheet of fiberglass can hold over 250 pounds per square inch. Using a hydraulic press, the spare fiberglass section and the acrylic section was subjected to forces simulating the expected loads during motor thrust. To do this, a steel plate was placed on top of and below the tube and pressed in the center. The coupler,


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fiberglass airframe, and acrylic airframe were all tested and each withstood the maximum weight of 1,500 pounds from the scale with no signs of wear or damage.

The fins for the final vehicle are twice as thick as the fiberglass bulkhead. Thus, the shear strength of a fin is greater than 250 pounds per square inch, the minimum tested strength of the bulkhead.

Epoxy

To test the Proline 4500 epoxy, a bulkhead was epoxied into an airframe, filleting one side to simulate how the bulkheads are incorporated in the full scale rocket. A 2x3 inch block was placed on the bulkhead to simulate the mounting hardware for the recovery system. Then, using the hydraulic press, pressure was applied to the bulk head. The force on the bulkhead reached 1,500 pounds and held this force for 60 seconds before it was released with no sign of wear or damage. To test the epoxy for mounting, the fins were mounted to a tube in the same manner as the prototype build. This will also replicate the fin mounting in the final build. The tube was then secured using clamps, and the hydraulic press was used to apply weight at the point of the fin furthest from the rocket. With 110 pounds of force applied at 5 inches from the airframe, the press provided enough torque to fracture the epoxy bond at the motor tube and the bond from the fins to the external airframe surface. Using τ = F x d, a torque of 550 inch-pounds is the maximum force applicable before the epoxy is compromised. After the CDR it became necessary to trim the fins by one inch to obtain better stability witch changes the maximum amount of perpendicular force on the fin before the epoxy is compromised. Using 550 in-pounds as the torque, and 4 inches as the distance, the maximum amount of force perpendicular to the fin is 137.5 pounds.

Workmanship

The mission success criterion provides key goals that must be met in order for the mission to be deemed successful. Completing these goals reflects directly upon the degree of workmanship of the vehicle design. The team approaches workmanship by understanding the crucial importance of building the vehicle as closely to the intended design as possible. The attachment, construction, fabrication, manufacture, and assembly of all structural elements dictate the overall robustness of the vehicle design. A primary concern is building the launch vehicle such that it has a safe and stable flight. It must possess an acceptable degree of survivability so that it may be reusable on the day of the official launch. The team understands that the vehicle is only as good as its construction. Proper care and attention must be taken in the construction of the vehicle. The team benefits from around the clock access to manufacturing facilities as well as a remote testing site.


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Recently, the team has been invited to produce precision components at the Polen facility in Granbury, TX. This facility provides aircraft quality precision tools including: analog calipers accurate to .001 inch, a manual lathe, a mill digitally measured to within .0001 inches, a CNC mill, PTC Creo Parametric, Solid Works, a band saw, a grinding station, a hydraulic press, and an extensive assortment of hand tools. The precision manufacturing process is overseen by facility owner Richard Keyt or Polen employees.

Vehicle Safety and Failure Analysis

Please refer to the Safety and Environment (Vehicle) section for the vehicle safety and failure analysis.

Full-scale Launch Results

March 14, 2013

Changes made to design:

None Changes made to procedures:

None

Date March 14, 2013

Location Hunewell Ranch, 32.2N, -98.2E, 1309MSL

Launch Rod 10 foot, button

Wind Speed 15 mph

Wind Direction SSW

Temperature 75° F

Pressure 30.21 inHg

Table 5: Launch Conditions


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Figure 23: Simulation Data

Figure 24: OpenRocket Rendering of Vehicle

Figure 25: Simulated Flight Data


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Stage Time Velocity Height AGL Lateral Displacement

Off-Rail 0.290 s 76.766 fps 10 ft 0 ft

Drogue Deployment

17.708 s 3.493 fps 4925 ft 627 ft

Main Deployment

64.471 s 92.478 fps 676 ft 135 ft

Impact 119.550 s 11.323 fps 0 ft 1377 ft

Component Apogee Time to Apogee Time of Flight

Drogue Raven 3

Corrupted Data Corrupted Data Corrupted Data

Drogue Stratologger

4947ft AGL 18.40 s 97.30 s

Main Raven 3

Corrupted Data Corrupted Data Corrupted Data

Main Stratologger

4935ft AGL 18.95 s Corrupted Data

GPS* 4977ft AGL 45.00 s 120.00 s

*The Garmin Astro DC40, which is secured inside the nose cone, reported a landing distance of 1,056 feet from the launch rail. The simulation predicted an apogee of 4,925 feet AGL, the average recorded apogee was 4,953 feet AGL, which is within 28 feet. A mass simulator was used to represent the weight of the final SMD payload. The vehicle was fully ballasted to 10% of the total vehicle weight, resulting in a total weight of 39.5 pounds. This was the first flight with the final vehicle design and is used for the full scale demonstration flight. The flight was a success and the vehicle was safely recovered and reusable at landing. The drogue Stratologger (secondary) flight data is shown in Figure 26.

Table 6: Simulated Flight Data

Table 7: Recovery Electronics Data


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The main Stratologger (primary) flight data is shown in Figure 27.

Figure 26: Drogue Stratologger Flight Data

Figure 27: Main Stratologger Flight Data


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March 15, 2013

Changes made to design:

None Changes made to procedures:

None

Date 15 March 2013

Location Hunewell Ranch, 32.2N, -98.2E, 1309MSL

Launch Rod 10 foot, button

Wind Speed 19 mph

Wind Direction SSW

Temperature 80 F

Pressure 33.2 inHg

Table 8: Launch Conditions

Figure 28: Simulation Data

Figure 29: OpenRocket Rendering of Vehicle


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Stage Time Velocity Height AGL Lateral Displacement

Off-Rail 0.280 s 78.994 fps 10 ft 0 ft

Drogue Deployment

18.202 s 1.501 fps 5306 ft 517 ft

Main Deployment

70.344 s 89.326 fps 690 ft 242 ft

Impact 128.980 s 11.120 fps 0 ft 1299 ft

Component Apogee Time to Apogee Time of Flight

Drogue Raven 3

5255ft AGL 18.96s 72.68s

Drogue Stratologger

5282ft AGL 19.60s 73.10s

Main Raven 3

5207ft AGL 18.72s 72.68s

Main Stratologger

5269ft AGL 19.65s 74.90s

GPS* Not Reliable Not Reliable Not Reliable

*The Garmin Astro DC40 reported a landing distance of 681ft from the launch rail. While the simulation predicted an apogee of 5306ft AGL, the average recorded apogee was 5253ft AGL, within 27ft of the desired apogee of one mile. The ballast was optimized for the purpose of reaching an apogee of one mile, at 1.3 pounds. The SMD payload was on board, and while the rocket did become treed, it was recovered. The SMD payload maintained telemetry throughout the flight and after landing. Due to the amount of time that was taken to recover the rocket from the tree, it was confirmed that recovery and payload electronics can maintain power for much longer than one hour.

Figure 30: Simulated Flight Data

Table 9: Simulated Flight Data

Table 10: Recovery Electronics Data


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For a more detailed analysis of the SMD flight, please refer to the Payload Criteria Section. The drogue Stratologger data is shown in Figure 31.

Drogue Raven3 is shown in Figure 32.

Figure 31: Drogue Stratologger Data

Figure 32: Drogue Raven3 Flight Data


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Main Stratologger is shown in Figure 33.

Main Raven3 flight data is shown in Figure 34.

Figure 33: Main Stratologger Flight Data

Figure 34: Main Raven3 Flight Data


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Mass Report

The unballasted mass summary of the launch vehicle and its three main sections are located in Table 11. Each vehicle section is also broken down into its individual components in Table 12. All listed mass calculations for the vehicle, sections, and individual components were obtained through the weighing of the components on two separate scales. All items weighing less than 2.2045 pounds were weighed on a JS-1000V digital scale with an accuracy of 0.0002 pounds. For heavier components, a Tanita 1583 baby scale was used. The baby scale has a weight capacity of 40 pounds with a graduation of 0.5 ounces for up to 20 pounds and one ounce for 20 to 40 pounds. The final launch vehicle has an unballasted mass of 35.9 pounds on-the-rail and a fully ballasted mass of up to 39.5 pounds. The variability of up to 10 percent of the vehicle mass for the ballast system allows for greater control of altitude for a wider range of atmospheric conditions. It was also found during testing that the mass of each motor varies slightly from the listed amount and is accounted for in each simulation with a variable mass component.

Mass Summary Section Mass (oz) Mass (lb)

Upper Body 148.99 9.31

Payload Housing 186.69 11.67

Booster 238.22 14.89

Total Mass (Launch) 573.90 35.87

Total Mass (Apogee) 570.03 32.00

Mass per Section Upper Body Component Mass (oz) Mass (lb)

Ballast System Bulkhead 4.85 0.303

Ballast System/Upper Eye Bolt1 2.54 0.159

Dog Tracker GPS 4.80 0.300

Main Deployment System 64.0 4.00

Main Ejection Charge 1 0.192 0.012

Main Ejection Charge 2 0.208 0.013

Main Shock Cord 9.30 0.581

Nose Cone 16.5 1.03

Upper Body Airframe 45.1 2.82

Upper Dog Barf 1.50 0.0940

Subtotal 148.99 9.31

Payload Housing Component Mass (oz) Mass (lb)

Acrylic Airframe 40.96 2.56

Drogue Avionics 5.952 0.372

Table 11: Mass Summary


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Drogue Avionics Bay Lid 11.696 0.731

Lower Coupler 22.4 1.4

Lower Payload Bulkhead 10.352 0.647

Main Avionics 5.952 0.372

Main Avionics Bay Lid 11.696 0.731

Upper Coupler 24.496 1.531

Upper Payload Bulkhead 11.744 0.734

USLI Electronics Payload 41.44 2.59

Subtotal 186.688 11.668

Booster Component Mass (oz) Mass (lb)

Centering Rings – 4 12.864 0.804

Drogue Ejection Charge 1 0.08 0.005


Drogue Parachute 3.04 0.19

Drogue Shock Cord 4.624 0.289

Fins – 4 21.6 1.35

Lower Body Airframe 44 2.75

Lower Dog Barf 1.056 0.066

Lower Eye Bolt 2.864 0.179

Motor Retaining Ring 4.96 0.31

Motor Tube 25.12 1.57

Motor2 3 117.92 7.37

Subtotal 238.224 14.889

Centering Rings – 4 12.864 0.804


1Mass of Ballast varies with configuration. 2Mass listed is for launch. The empty mass is: 3Mass variation found in each motor.

Recovery Subsystem

Design Defense

The final recovery system design has been tested with sub-scale prototype launches as well as two full-scale launches.

Table 12: Mass by Subsection


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The parachute systems are as robust as the shock harnesses. In the upper body airframe, a layer of 1.5 ounces of dog barf wadding, 33 feet of z-folded 0.375 inch tubular Kevlar line, and a Kevlar deployment bag protect the main parachute from the heat of the ejection charge. No heat damage is visible on the main parachute, pilot parachute, or shroud lines due to test launches. Slight singeing is visible on the front flap of the Kevlar bag, but its integrity has not been compromised. Similarly, in the booster section, a layer of 1.05 ounces of dog barf wadding, 16 feet of z-folded 0.375 inch tubular Kevlar line, and a Nomex blanket protect the drogue parachute from the heat of the ejection charge. No heat damage is visible on any of the drogue parachute components. Prior to each launch, the recovery team inspects the shock cord for any fraying or kinks. All kinks are removed before the line is z-folded and secured with either painter's tape or a thin rubber band. Fraying has not yet occurred to warrant a new line, but it is planned to fly with fresh harnesses on launch day. The manufacturer rating for the 0.375 inch tubular Kevlar from Rocketry Warehouse used in all test flights is 3,600+ pound-force. Since the design is shedding weight, the highest load experienced on launch day is expected to be less than 1,750 pound-force, the highest test load experienced, as analyzed from Featherweight Raven 3 data captured from the flight on February 19, 2013. This gives a worst-case-scenario factor of safety rating (FOS) of 3600/1750 = 2.057, which exceeds the 2.000 required. The highest drogue deployment velocity experienced on flights with the 0.375 inch shock line used in all test flights occurred on January 26, 2013. After burnout, the vehicle weighed slightly less than 30 pounds, and experienced 11.68 G on drogue parachute deployment. This resulted in a maximum shock of 350.4 pound-force, for a worst-case-scenario FOS of 3600/350.4 = 10.274. The only knots in the 40 foot shock harness are at the connection point to the nose cone, and at the connection point to the main avionics bay coupler. The knots in the drogue (20 foot) shock harness are located at the connection point to the motor casing, the connection point to the drogue avionics coupler, and four feet from the connection point to the drogue avionics coupler. This latter knot serves as an attachment point for the drogue parachute. All knots are figure-eight knots. No adhesives are used to secure any portion of either harness. Although the line is heat resistant Kevlar, each harness is shielded from the blast of the initial ejection charge firing by dog barf. It is recognized that the secondary charge firing could cause heat damage to the shock cord. All welded eyebolts have remained intact after each flight. No attachment hardware including quick links, washers, lock washers, and all-thread, has been damaged as the result of any test flight.


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Power loss to the altimeters during flight can result due to the battery coming temporarily disconnected from the shock force of main parachute deployment. While the former primary altimeters, the Featherweight Raven3s have shut down and reentered flight mode due to this power loss, the former secondary altimeters, the Perfectflite Stratologger SL100s, remain in flight mode and gathering data through the surge. All four recovery altimeters are now Perfectflite Stratologger SL100s. The Featherweight magnetic arming switches have not malfunctioned to this point. Externally visible LED lights wired through each magnetic switch power on only when the switch is armed by the magnet. These lights confirm power to the recovery altimeters. The Garmin Astro DC-40 transmitter in the nose cone is powered on prior to packing the main parachute components in the upper airframe. Four screws with plastic anchors hold the nose cone, where the transmitter and the ballast system are housed to the upper airframe.

Structural Elements

In order to enable deceleration of the three tethered sections on descent, 0.375 inch tubular Kevlar shock cords have been implemented based on strength and flame-proof construction. The length of the shock cord for the drogue parachute is 20 feet. Based on the advice of the team mentor, the main parachute shock cord should be twice as long as the drogue parachute shock cord. The main parachute shock cord is, in effect, 40 feet long.

The shock cord for each parachute is z-folded and bound with rubber bands. This reduces the risk of entanglement and saves space in the parachute compartments. The harnesses are attached to the vehicle via stainless steel quick links through a figure-eight knot in the end of the line. This connects to welded eyebolts on airframe couplers. The motor casing within the booster section has a welded eyebolt installed. This eyebolt is used to attach the booster section to the 20 foot shock cord. A section of shock cord epoxied to the interior wall of the booster section in case of failure of the eyebolt on the motor housing is attached to the drogue parachute shock cord. The other end of the shock cord is attached to a welded eyebolt installed on the exterior bulkhead of the drogue avionics bay. This effectively tethers the booster section to the payload housing structure. The drogue parachute shroud lines are attached to a figure-eight knot in the shroud line four feet below the drogue avionics bay. The exterior bulkhead of the main avionics bay also has an eyebolt installed. This eyebolt is used to attach the 40 foot shock cord to the payload housing section. The other end of this line is attached to a welded eyebolt on the coupler between the nose cone and the upper body airframe. The main parachute shroud lines are also anchored here via a swivel between the four lines and the quick link.


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A „d‟ shaped stainless steel ring at the apex of the main parachute is attached via quick link to a square stainless steel ring on the end of a nylon strap on the inside of the deployment bag. A quick link is also used to attach a pilot drogue parachute to a square stainless-steel ring on the end of a nylon strap on the outside of the deployment bag. The shroud lines are woven through elastic straps on the front of this bag.

Electrical Elements

In order to stage dual deployment, two sets of altimeters are implemented. Each system consists of two Perfectflite Stratologger SL100 altimeters mounted to a 0.19 inch thick, 5 inch wide, 8.125 inch long plywood sled and the power supply. The magnetic arming switch is mounted on the opposing side from the altimeters and sits at the edge of the inner avionics bay wall.

A nine-volt battery supplies power to each set of altimeters through a Featherweight Magnetic Arming Switch. Each switch is able to be armed externally using a magnet provided by the manufacturer. Based on trial and error, the range of the magnet for arming the switch is best at one inch from the exterior of the airframe. It is confirmed that the switch has been armed and is supplying power to the altimeters by an externally visible LED light. This light remains on until the switch is disarmed and power is no longer being supplied to the altimeters. Each altimeter is wired to a black powder ejection charge. The charges are each made from one J-Tek 10 electric match cut to three feet for the booster section and four feet for the nose cone section. The pyrogen-coated end of the electric match rests at the bottom of a 13millimeter by 100millimeter Polystyrene test tube, with the leads pushed through a 0.125-inch hole drilled into the bottom. The electric match is secured to the tube by hot glue on the exterior of the hole. The exposed leads are secured to a male crimp connector. The charge leads from each altimeter are secured to a female crimp connector on the opposite side of the avionics coupler. Each electric match of a certain black powder charge weight is connected to the appropriate altimeter terminals. The charge canisters are placed horizontally at the bottom of the appropriate airframe section. The leads are secured with a piece of Gorilla tape to ensure that the crimp connection is not compromised.

Redundancy Features

The recovery system employs a dually redundant ejection charge system to ensure airframe separation and parachute ejection at the correct flight stage. Two altimeters of the same make and model are programmed to fire separately. The primary will fire a


32


charge of the calculated size for the compartment at the desired time, while the secondary will fire a slightly larger charge at a delayed time. In the drogue parachute avionics bay, the primary SL100 is programmed to fire at barometric detection of apogee. A 2.6 gram FFFg black powder charge is to be ignited and cause separation of the booster section from the forward airframe. The secondary SL100 is programmed to fire at one second past barometric detection of apogee. A 2.8 gram FFFg black powder charge is to be ignited in case of failure of the primary charge to ensure separation and ejection of the drogue parachute.

In the main parachute avionics bay, the primary SL100 is programmed to fire at 700 feet above ground level (AGL). The secondary SL100 is programmed to fire at 650 feet AGL. A time delay is not used for main parachute ejection, since the time delay feature on these altimeters is only available for apogee. However, the 50 foot delay should be sufficient to ensure that both altimeters in the main avionics bay do not fire simultaneously.

The manufacturer-reported calibration accuracy of these altimeters is within 0.05%. This corresponds to an error of 0.35 feet at 700 feet AGL between two typical SL100 altimeters in the main parachute avionics bay, supposing they entered flight mode at the same time. Furthermore, the manufacturer-reported measurement precision of these altimeters is 0.1% of the reading, plus a foot. This corresponds to an error of 1.7 feet at 700 feet AGL between two typical SL100 altimeters in the main parachute avionics bay, supposing they entered flight mode at the same time. A difference of 2.05 feet between the main avionics bay altimeters is typical, which is well within the 50 foot delay employed.

As mentioned in the previous section on structural elements, the recovery system also features redundant tethering of the booster section. A 24 inch section of 0.375 inch tubular Kevlar with a loop at the forward end is epoxied to the interior wall. The drogue parachute shock cord is attached to this loop via quick link, then to the eyebolt on the motor housing. This serves as a backup connection of the drogue parachute harness to the booster section, ensuring tethering of the booster section to the forward airframe in the event of failure of the eyebolt on the motor housing.

Parachutes

The final recovery design consists of the ejection of a drogue parachute at apogee, followed by the ejection of a main parachute at a height of between 700 and 650 feet AGL on descent. The drogue parachute is protected from ejection charge firing by a Nomex cloth and dog barf wadding. The main parachute is also protected from ejection charge firing by dog barf. Additionally, the main parachute is packed into a Kevlar TAC-9B Deployment Bag to reduce the risk of entanglement of the shroud lines on ejection. To further ensure deployment of the main parachute, a pilot drogue parachute is attached to the bottom of the deployment bag to help pull the bag out of the airframe on


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separation. The deployment bag is shown in Figure 35. The operation of the deployment bag is specified in Figure 36.

Figure 35: Packed Deployment Bag

Figure 36: Deployment Bag Operation


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As in the CDR, the choice of drogue parachutes remains the Sky Angle CERT-3 24", with a drag coefficient of 1.16. Likewise, the choice of main parachute remains the Sky Angle CERT-3 XX Large 120", with a drag coefficient of 2.92. Each parachute is cross-form and constructed of 1.9 ounce silicon-coated rip-stop Nylon. All seams are reinforced with Nylon webbing to reduce the probability of shroud line disconnection on deployment. With four shroud lines, each made of 5/8 inch woven tubular nylon for durability, these parachutes have a low probability of entanglement. The length of the shroud lines on the main parachute is ten feet, while that on each drogue parachute is two feet. Further reducing the risk of entanglement upon deployment is the 7,000 pound tested swivel between the main parachute shroud lines and shock cord. The total launch vehicle weight is 35.9 pounds, but the team has tested the maximum (10%) ballast for a launch vehicle weight of up to 39.49 pounds. At apogee, burnout will have occurred already, to give a vehicle weight of up to 35.62 pounds. 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.

√

From the above calculations, it can be concluded that the launch vehicle will experience a combined impact force of less than 75 foot-pounds should it land at a speed of no more than 11.645 feet per second. This result is used below to calculate the minimum diameter of the main parachute. Assumptions in the calculation are that the vehicle is descending at a constant speed, and that the downward motion is simple, along the z-axis.


35


( )

√

√

The minimum diameter of the main parachute is thus 2a = 9.918ft. Based on commercial availability of parachutes, a 10 foot diameter parachute theoretically suffices for the main. The drag coefficient of 2.92 is provided by the manufacturer for the chosen and tested main parachute. The calculation for the descent rate of the vehicle under the 10 foot main parachute follows.

√

√

The calculated descent rate of the vehicle after main parachute deployment at 700 feet AGL is thus 11.645 feet per second. With this figure, it is possible to calculate the total drift of the vehicle after main parachute deployment until landing. The drift of the vehicle from main parachute deployment until landing follows, assuming a 15 mile per hour horizontal wind.

With a fifteen mile per hour wind, the launch vehicle will have drifted approximately 1,322 feet from the launch pad under the main parachute before landing. Since the maximum allowable drift is 2,500 feet in a fifteen mile per hour wind, the drift between drogue and main parachute deployment can be at most 1,178 feet. Using this maximum drift between drogue and main parachute deployment, the minimum descent rate for the


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vehicle under the drogue parachute is calculated, assuming an apogee of one mile, or 5280 feet AGL.

Thus the minimum descent rate from apogee to main parachute deployment to ensure a total drift of no more than 2,500 feet by landing is 85.535 feet per second. With this figure, it is possible to determine the appropriate size of the drogue parachute. The maximum diameter of the drogue parachute that would allow for this descent rate from apogee is determined below. Assumptions made in this calculation are that the vehicle is descending at a constant rate, and that this downward motion is simple, lying solely along the z-axis.

( )

√

√

The minimum diameter of the drogue parachute is thus 2a = 2.142ft. Based on commercial availability of parachutes, a two foot diameter parachute theoretically suffices for the drogue. The drag coefficient of 1.16 is provided by the manufacturer for the chosen and tested drogue parachute.

Drawings and Schematics

Electrical


37


The wiring configuration of the altimeters is given in Figure 37.

Figure 38 shows the Stratologger Software flow diagram.

Figure 37: Altimeter Wiring Configuration


38


Figure 39 shows the GPS software flow diagram.

Figure 38: Stratologger Software Flow Diagram


39


Figure 39: GPS Software Flow Diagram


40


Structural

The following drawings are of the structural components of the avionics bay. The avionics bay are shown below in Figure 40.

The features of the attachment bulkhead within the avionics coupler is shown below in Figure 41.

Figure 40: Avionics Bay Structural Components

Figure 41: Attachment Bulkhead Inside Avionics Coupler


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The dimensions of the couplers that house the avionics bays are shown in Figure 42.

Figure 43 contains an exploded view for the structural assembly of the avionics bay.

Recovery Tracking

The recovery system includes a GPS transmitter and a handheld receiver, as seen in Figure 44. The transmitter, as stated in the CDR, is the Garmin Astro DC-40, which runs up to 48 hours on a rechargeable lithium-ion AA battery. This unit is attached to the all-thread support structure of the ballast system in the nose cone.

Figure 42: Couplers in Profile

Figure 43: Structural Assembly of the Avionics Bay


42


The handheld receiver is the Garmin Astro 320, which runs on two AA batteries for up to 20 hours. This unit is located at the ground station with a member of the recovery team. The range of transmission accuracy between the DC-40 and the 320 is reported by the manufacturer to be up to nine miles in clear, unobstructed terrain.

The output power of the DC-40 is two Watts. Use of any high-powered radio near the 320 could cause irreparable damage to the unit. Note that a high-powered radio is defined as having a power output of at least five Watts.

Transmitting frequencies picked up by the receiver are in the Multi-Use Radio Service (MURS) band. The MURS band consists of 151.8200MHz, 151.8800MHz, 151.9400MHz, 154.5700MHz, and 154.6000MHz. The default setting on the 320 is 151.8200MHz, but any of the five MURS frequencies may be chosen.

The recovery team conducted a series of range tests. These included leaving the DC-40 in the rocketry lab at the university math building and driving off with the 320 until the signal was lost, and leaving the DC-40 at the launch site on Hunewell Ranch and driving or walking off until the signal was lost. It was found that in obstructed terrain, including neighborhoods, 0.792 miles (4181.760 feet) was the shortest range of accuracy. Thus the system should be sufficient to locate the vehicle within the 2500 foot allowable landing radius on launch day.

System Sensitivity

Aluminum flashing (0.01 inches in thickness) lines all surfaces of each avionics bay. Since aluminum is a conductor of electricity, this forms a Faraday cage around the recovery altimeters. Thus, not only are the avionics bays shielded from any onboard or nearby radio frequencies reflected by the flashing, but also from magnetic waves generated by onboard electronics.

Figure 44: GPS Device and Wireless Remote


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If there is a sufficiently strong source of magnetic waves nearby, it is possible that the switch could be armed. This would switch the LED(s) of the armed avionics bay(s) on, however, enabling the team to disarm the affected system(s). Furthermore, the altimeters will not send voltage to an ejection charge unless they have entered flight mode, so inadvertent excitation of the recovery altimeters can be safely mitigated through deactivation. As discussed in the previous section on recovery tracking, the recovery tracking system is susceptible to high-power radio devices. None of the onboard radio transmitters has an output power of greater than two Watts. This can be verified in the description of the SMD payload.

Parachute Justification

As calculated previously, the most appropriate sizing of the main and drogue parachutes is 10 and two feet, respectively. These calculations take into account deployment process and vehicle mass. The attachment scheme of the main parachute comes directly from the manufacturer. This scheme was validated on the last three test flights. The attachment scheme for the drogue parachute, as described in the previous sections on structural elements and redundancy features, has been optimized through test flights. As documented in the CDR, one static ejection test has been completed based upon calculated black powder weights. The ejection charge sizing is appropriate for the length of the airframe section for each parachute. This has been thoroughly justified through test flights, as charge sizing has enabled separation upon being prepared properly.

Safety and Failure Analysis

Following is a log of the seven test flights completed since the CDR. For each flight, failure modes are identified. Each failure mode is then analyzed in terms of components affected, components responsible, potential causes, and potential hazards. Mitigations of these hazards are then identified and enacted on subsequent flights.

The first five test flights were conducted using the full-scale Prototype 1, which was built to specifications from the CDR and gradually modified to incorporate many of the specifications in the FRR. The last two, namely the full scale demonstration flight (3/14/2013) along with a full scale test flight with SMD payload (3/15/2013) were conducted using the finalized design, which was built to specifications in the FRR. These last two flights reflect expected performance on competition launch day.


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Test Date: January 26, 2013

Brief Description of Test: A dual deployment test launch was conducted using the modified full-scale Prototype 1 rocket at Hunewell ranch around noon. The motor used was a Cesaroni L585. For main parachute ejection the altimeter used was a Perfectflite Stratologger SL100, equipped with 5.8g packed 3F black powder. A Stratologger was also used for backup ejection of the drogue parachute, with a 2.8g packed 3F black powder charge. The primary altimeter used for drogue parachute ejection was a Featherweight Raven 3, with a 2.6g packed 3F black powder charge. It was shown in this flight that the attachment scheme for the main parachute deployment bag is sufficient for successful deployment of the main parachute with minimal entanglement. Each parachute was ejected fully and as programmed. The launch that vehicle was recovered in a mostly clear area. The only damage sustained to any component was that potentially incurred to the Garmin Astro DC40 when it was ejected from its attachment point.

Failure Mode

Component(s) Responsible

Component(s) Affected

Potential Cause(s)

Potential Hazard(s)

Mitigation

Untethered transmitter unit

Attachment to drogue parachute shock cord

Direct location of landed launch vehicle

Tape weakened by heat of ejection charge and broke on force of impact

Lost transmitter makes it harder to find rocket, could be disqualified if this happens in competition

New attachment scheme involving zip ties and back plate to secure DC40 to drogue shock cord

Table 13: Test Flight 1/26/13


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Test Date: February 14, 2013

Brief Description of Test: A dual deployment test launch was conducted using the modified full-scale Prototype 1 rocket at Hunewell ranch at 3 p.m. The motor used was a Cesaroni L585. For main parachute ejection the altimeters used were a Perfectflite Stratologger SL100 programmed to fire at 650ft AGL on descent, equipped with a 5.5g packed 3F black powder charge, and a Featherweight Raven3 programmed to fire at 704ft AGL on descent, equipped with a 5.0g packed 3F black powder charge. A Stratologger was used for backup ejection of the drogue parachute at 2s past apogee, with a 2.6g packed 3F black powder charge. The primary altimeter used for drogue parachute ejection was a Featherweight Raven 3, with a 2.9g packed 3F black powder charge. Also on board was a small portion of the payload for data gathering. The Garmin Astro DC40 GPS unit was secured inside the nose cone rather than to the drogue parachute shock cord as in the last flight. It was shown in this flight that the main parachute shock cord should be located between the ejection charges and the deployment bag, rather than between the deployment bag and the avionics coupler, in order to prevent entanglement. It was also determined that ribbon wire would be used for all avionics connections to reduce the risk of disconnection before landing.

Failure Mode



Potential Cause(s)

Potential Hazard(s)

Mitigation

Main Parachute not Envelope

Main parachute shock cord

Main parachute shroud lines did not release, causing main parachute to not be released from bag until impact

Placement between bag and coupler caused shroud lines to push into and tangle with shock cord upon ejection

Ballistic impact, damage to vehicle or on-board electronics

Place main parachute shock cord between bag and charges



46


Test Date: February 19, 2013

Brief Description of Test: A dual deployment test launch was conducted using the modified full-scale Prototype 1 rocket at Hunewell ranch at 2 p.m. The motor used was a Cesaroni L585. For main parachute ejection the altimeters used were a Perfectflite Stratologger SL100 programmed to fire at 650ft AGL on descent, equipped with a 5.6g packed 3F black powder charge, and a Featherweight Raven3 programmed to fire at 704ft AGL on descent, equipped with a 5.4g packed 3F black powder charge. A Stratologger was used for backup ejection of the drogue parachute at 2s past apogee, with a 3.4g packed 3F black powder charge. The primary altimeter used for drogue parachute ejection was a Featherweight Raven 3, with a 3.2g packed 3F black powder charge. No payload was on board. It was shown in this flight that the new placement of the main parachute shock cord was successful, and that rubber grommets should be used to cushion the attachment points of the acrylic payload section to the avionics couplers.

Failure Mode



Potential Cause(s)

Potential Hazard(s)

Mitigation

Reusability of acrylic payload section

No shock absorption between acrylic and screws

Integrity of acrylic at attachment points to drogue avionics coupler

Force of impact

Damage to payload electronics, reusability of system

Rubber grommets to cushion acrylic from screws attaching it to the avionics couplers



47


Test Date: March 7, 2013

Brief Description of Test: A dual deployment test launch was conducted using the modified full-scale Prototype 1 rocket at Hunewell ranch at 2p.m. The motor used was a Cesaroni K510. For main parachute ejection the altimeters used were a Perfectflite Stratologger SL100 programmed to fire at 650ft AGL on descent, equipped with a 5.6g packed 3F black powder charge, and a Featherweight Raven3 programmed to fire at 704ft AGL on descent, equipped with a 4.9g packed 3F black powder charge. A Stratologger was used for backup ejection of the drogue parachute at 2s past apogee, with a 2.9 packed 3F black powder charge. The primary altimeter used for drogue parachute ejection was a Featherweight Raven 3, with a 2.6g packed 3F black powder charge. No payload was on board. It was shown in this flight that the new deployment bag works, but that the charge leads need to be cut to fit.

Failure Mode



Potential Cause(s)

Potential Hazard(s)

Mitigation

Ignition of Primary Drogue E-Match

Single-core ground wire

Ground wire pulled out

Leads were too long and became tangled with drogue parachute shroud lines

Failed drogue parachute ejection if secondary charge doesn't ignite, high-speed main parachute ejection

Cut e-match leads to fit tube length loosely, try to find braided wire connectors

Integrity of Drogue Parachute

Length of drogue e-match leads

Reusability of parachute

Entanglement of e-match leads with drogue parachute shroud lines

Drogue parachute not reusable, high-speed main parachute ejection

Cut e-match leads to fit tube length loosely



48



Brief Description of Test: A dual deployment test launch was conducted using the modified full-scale Prototype 1 rocket at Hunewell ranch at 10a.m. The motor used was a Cesaroni K555. For main parachute ejection the altimeters used were a Perfectflite Stratologger SL100 programmed to fire at 650ft AGL on descent, equipped with a 3.2g packed 3F black powder charge, and a Featherweight Raven3 programmed to fire at 704ft AGL on descent, equipped with a 2.6g packed 3F black powder charge. A Stratologger was used for backup ejection of the drogue parachute at 2s past apogee, with a 5.1g packed 3F black powder charge. The primary altimeter used for drogue parachute ejection was a Featherweight Raven 3, with a 4.6g packed 3F black powder charge. No payload was on board. It was shown in this flight that all eyebolts must be welded for secure connection to be maintained. Further, it was decided after collecting the altimeter data from this flight that all four recovery altimeters will be Perfectflite Stratologger SL100's since the temporary power loss experienced upon airframe separation does not affect flight mode.

Failure Mode



Potential Cause(s)

Potential Hazard(s)

Mitigation

Vehicle remaining tethered

Drogue avionics coupler eyebolt failure

Altimeter power loss, no main parachute deployment, GPS transmitter damage, acrylic payload section shattered, one fin cracked, one pulled eye bolt, irreparable body damage

Eyebolt not welded, caused untethering of drogue harness at apogee

Vehicle not reusable, vehicle not locatable, damage to onboard electronics

Switch all four recovery altimeters to Perfectflite Stratologger SL100's, weld all eyebolts



49



Brief Description of Test: A dual deployment test launch was conducted using the fully-ballasted (10%) full-scale Prototype 2 rocket at Hunewell ranch at 5:45p.m. The motor used was a Cesaroni L1720. For main parachute ejection the altimeters used were a Perfectflite Stratologger(secondary) SL100 programmed to fire at 650ft AGL on descent, equipped with a 5.2g packed 3F black powder charge, and a Featherweight Raven3(primary) programmed to fire at 704ft AGL on descent, equipped with a 4.8g packed 3F black powder charge. A Stratologger(secondary) was used for backup ejection of the drogue parachute, with a 2.8g packed 3F black powder charge. The primary altimeter used for drogue parachute ejection was a Featherweight Raven3(primary), with a 2.6g packed 3F black powder charge. A mass simulator of the payload was onboard. This flight was successful.

Failure Mode



Potential Cause(s)

Potential Hazard(s)

Mitigation

None None None None None None



50


Test Date: 15 March 2013

Brief Description of Test: A dual deployment test launch was conducted using the full-scale Prototype 2 rocket at Hunewell ranch at 6:45p.m. The motor used was a Cesaroni L1720. For main parachute ejection the altimeters used were a Perfectflite Stratologger(backup) SL100 programmed to fire at 650ft AGL on descent, equipped with a 5.2g packed 3F black powder charge, and a Featherweight Raven3(main) programmed to fire at 704ft AGL on descent, equipped with a 4.8g packed 3F black powder charge. A Stratologger(backup) was used for backup ejection of the drogue parachute at 1s after apogee, with a 2.8g packed 3F black powder charge. The primary altimeter used for drogue parachute ejection was a Featherweight Raven 3(main), with a 2.6g packed 3F black powder charge. The SMD payload was on board. The ballast system was optimized for the purpose of hitting an apogee of one mile, at 1.3 pounds. The results of this flight demonstrated reliability of the simulation.

Failure Mode



Potential Cause(s)

Potential Hazard(s)

Mitigation

Main Parachute became entangled

Pilot drogue for deployment bad

Minor acrylic damage from impact with tree

Shock chord entanglement

Very high descent velocity

Verify attachment and packing scheme

Mission Performance Predictions

Mission Performance Criterion

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. After apogee and descent, the entire vehicle lands within 2,500 feet of the launch pad. In addition to meeting the mission success criteria, a perfect launch will reach apogee at exactly 5,280 feet. With accurate simulation conditions, an appropriate ballast mast can be determined to achieve a perfect launch.

Flight Profile Simulations, Altitude Predictions, Weights, and Actual Motor Thrust Curve Flight simulations are performed on OpenRocket, an open-source software suite for rocketry. OpenRocket allows the user to completely construct the vehicle profile, define component materials and masses, select from commercially available motors, and simulate the entire flight based upon atmospheric conditions.



51


The flight simulation for the final design to be used on the official launch day is shown in Figure 45. With the vehicle profile, actual weights of all components, and the selected L1720-WT motor, the predicted altitude of the official flight is 5315 feet AGL with zero ballast weight in a 15 mile per hour wind. A summary of the input parameters for the simulation is given in Figure 46. These parameters reflect proposed conditions in Huntsville, AL. Careful attention was paid to ensure that simulation weights accurately reflect the actual weights of all components of the launch vehicle. These weights are completely itemized in the Mass Report section. The total vehicle mass with motor is 35.9 pounds unballasted. The overshoot of the target apogee can be reduced by implementing ballast weight. The most recent simulations suggest that with an added ballast weight of 4 ounces, totaling to 36.1 pounds with motor, the vehicle will reach apogee at approximately 5277 feet. The optimal vehicle design, for flight in 15 mph winds, will include the 4 ounces of ballast weight, creating a stability margin of 1.67. With a gross liftoff weight of 36.1 pounds (this includes total vehicle weight and 4 ounces of ballast), the thrust to weight ratio is 10.9:1 based upon the average thrust of the L1720 motor. The rail exit velocity is simulated to be 81.5 feet per second. During the flight, a maximum velocity of 680 feet per second (Mach 0.61) and maximum acceleration of 365 feet per second squared. Multiple trials with the same input parameters in OpenRocket show that the predicted apogee can vary by approximately plus or minus 10 feet, due to turbulence and non-constant wind speeds. Along with a 10% variability in the commercially available motors, these represent the most significant factors that are not in control by the team. Furthermore, these two factors will contribute greatest to error in the predicted versus actual performance.


52


Figure 45: Final Vehicle Simulation

Figure 46: Input Parameters for Final Simulation


53


As previously presented in the CDR, a static motor test was conducted to obtain an actual motor thrust curve that is provided in Figure 47. The actual motor thrust curve closely reflects the neutral thrust curve provided by the motor manufacturer in shape.

Thoroughness and Validity of Analysis, Drag Assessment, and Scale Modeling Results

As mentioned above, despite thorough detail in modeling the vehicle in OpenRocket, there still exists a level of uncertainty due to the uncontrollable factors of motor variability and wind turbulence variability. OpenRocket offers an approximation for the standard deviation of wind speed based on turbulence intensity; however, the calculations are based on simulated numbers. These variance values will be compared against data from the National Oceanic and Atmospheric Administration, as detailed in the Science Value section, to determine their accuracy. Upon verifying more precise values for wind speed standard deviation, the margin of error can be reduced and an apogee of 5,280 can be more easily achieved. Furthermore, the variability of the motor can only be accurately quantified by numerous thrust tests. Due to the price and availability of these motors, it is not feasible to burn multiple motors to determine variance between them. Cesaroni provides an average thrust curve for the L1720, but without a large amount of sample data these averages lend no insight towards the standard deviation.

Figure 47: L1720-WT Actual Thrust Curve


54


In lieu of the aforementioned tests and further analysis to determine any correlation between the variances, a large scale simulation analysis has been devised to observe overall interference from these factors. Using a custom-built automation macro, 944 simulations were conducted in OpenRocket to simulate the conditions of the official launch day in Huntsville, AL. The flights, as expected, differed significantly in their values and ranges across the sample. By conducting the experiment on such a large scale the unknown variances become easier to analyze via significant trends within the data pool. Using data extracted from the simulations confidence intervals were established to demonstrate the probability of achieving specific goals with the competition vehicle. As demonstrated in Table 20, different ranges of values offer different probabilities. According to the simulation analysis the launch vehicle has only a 5.5% chance of hitting exactly a mile, but as you expand the neighborhood of tolerance the probabilities rise significantly, resulting in a significantly high chance of near-mile apogee.

Apogee (feet) 5280 ± 5 ± 10 ± 15 ± 20

Probability (%) 5.530 51.374 83.624 96.328 99.465

Lateral Drift (feet) 1900 ± 50 ± 100 ± 150 ± 200

Probability (%) 0.534 49.909 82.174 95.653 99.290

The statistics suite RStudio was used to create the analysis script. The confidence intervals were built around a basic statistical analysis of the simulation sample. Upon finding the sample mean and sample variance, those values were used to build Probability Density Functions (PDF) to give a graphical representation of the likelihood of specific flight patterns. Figures 48 and 49 illustrate the functions for apogee and lateral distance. The range value depicts the chances of the variable in question taking corresponding values. Though the functions demonstrate relatively low probabilities of landing at any particular value, integrating this function across an interval greatly increases those probabilities. The table above reiterates this notion, that as the window of expected values widens, the chance of falling within that range grows exponentially.

Table 20 - Confidence Intervals for Apogee and Drift


55


As the team continues testing with full scale launches, the differences between actual performance and simulated performance of the vehicle and recovery system will become clearer. It is very important to understand and analyze these differences in order to make the simulation more accurate and valuable for altitude predictions. All parameters must be selected such that they most accurately represent the actual parameters of the vehicle flight. The utility of simulations as a performance measure is best illustrated by the successful launch of March 15, 2013. As the team began testing a full scale rocket with a finalized

Figure 48 - PDF of Apogee for Competition Flight

Figure 49 - PDF of Lateral Distance for Competition Flight


56


design, the discrepancies between real and virtual measurements dwindle. By exploring the many detailed settings of OpenRocket, the recovery team discovered means of introducing wind turbulence to the simulations. Though this presents some level of uncertainty, a non-constant wind model aligns more closely with real flight data. Since implementing these minor changes, the simulations have become much less erratic. Whereas simulations in earlier documents could have inaccuracies of up to 10%, the March 15, 2013 flight performed to within two feet of predictions. By emphasizing the verisimilitude of the virtual model, the suggested ballast weights bring the vehicle closer to a perfect flight on competition day. Drag assessment for the launch vehicle concerns two primary areas: vehicle profile characteristics and recovery system components. The vehicle experiences drag due to the shape and surface area of the nose cone as well as the surface area of the fins due to thickness. To date, no theoretical calculations of the drag profile of the launch vehicle have been completed. All predictions for the flight performance of the vehicle have been based on OpenRocket simulations. Lacking any facilities capable of testing or measuring aerodynamics, the team must currently rely on mathematical deduction and simulation software to determine drag characteristics. The recovery system creates drag from the drogue and main parachutes. Theoretical calculations have been completed to estimate landing radius of the vehicle due to parachute deployment events. These calculations are compared to the simulated landing radius from OpenRocket. Fortunately, weather cocking, which further reduces the actual landing radius of the vehicle, is taken into account by the OpenRocket simulation. This yields a smaller landing radius than theoretical calculations. Multiple sub-scale test launches were conducted to gain experience and understanding pertaining to the successful dual deployment of parachutes. All scale modeling results are from OpenRocket and provided above in the “Subscale Flight Results” section. Upon conducting further test launches, the goal is to minimize the difference between simulated performance and actual performance. Although the primary tool utilized for this purpose is a flight simulator, further modeling may be necessary to identify inconsistencies between simulated flight parameters and actual flight parameters. As discussed below, quantified differences between simulated and actual CG/CP calculations have been developed.

Stability Margin and the Actual CP and CG Relationships and Locations

The locations of the center of pressure (red) and center of gravity (blue) are shown in Figure 50. These are measured as distances from the tip of the nosecone. The resulting stability margin is 1.67 caliber, yielding a stable vehicle for flight.


57


The simulated measurements for center of gravity (CG) and center of pressure (CP) are:

An itemized mass budget was built to show the absolute locations, lengths, and weights of each component of the rocket. Software was developed to sum the itemized weights on either side of a predefined CG value. After several iterations of the software and refinements of the defined CG value, the program suggested a theoretical CG of 56.83 inches to balance the rocket. Using generalized equations from the 2012 NASA Advanced Rocketry Workshop (ARW) handbook these measurements were calculated by hand. According to the handbook, certain conditions must be satisfied to ensure the Barrowman equations adequately model the theoretical CP. These conditions are as follows:

The angle of attack (α) of the rocket is near zero (less than 10 degrees)

The speed of the rocket is much less than the speed of sound

The air flow over the rocket is smooth and does not change rapidly

The rocket is thin compared to its length

The nose of the rocket comes smoothly to a point

The rocket is an axially symmetrical rigid body

The fins are thin flat plates

Since each condition is met by the rocket, the Barrowman equations should accurately represent the theoretical CP. These calculations are as follows: Nose cone:

where L denotes length

Airframe: for all portions of the airframe

Figure 50: Center of Gravity and Center of Pressure


58


Fins:

[

]

[

( )

√ (

)

]

[

]

[

(

)

√ ( )

]

Where:

(all units are in inches)

Using the dimensions of the fin, as shown in Figure 15, the various chords were derived.

(

)

( )

(

)


59


Combined:

Table 21 compares the calculated and simulated values of CG and CP with measurements in inches. The small error values demonstrate that the simulated values are quite accurate. Incorporating these simulated values in further analysis of the rocket will introduce negligible errors.

Measurement Calculated Simulated Error

CG 56.80 in 56.83 in 0.053%

CP 67.11 in 66.03 in 1.64%

Stability Margin

1.87 cal 1.67 cal 11.98%

Kinetic Energy Management

The vehicle‟s kinetic energy for each independent section is calculated for the two phases of the mission. These are considered with the full vehicle weight including the four ounce ballast for optimal altitude performance. For each phase, a maximum descent velocity is used along with the mass of each section to determine kinetic energy of that section. The first consideration ends at main parachute deployment. At this point, the vehicle has already reached a terminal velocity of 83.14 feet per second under the drag of the drogue parachute. The second consideration is at landing when the vehicle has reached terminal velocity under the drag of the main parachute. At 9.95 feet per second upon landing, the kinetic energy of each vehicle section can be verified to be no more than 75 foot pound force at impact. The results of the kinetic energy calculations are given in Table 22. Note that the upper body airframe includes the nose cone and ballast.

Table 21: Calculated versus Simulated CG and CP Measurements


60


Phase Upper Body

Airframe (9.56 lbs)

Payload (11.67 lbs)

Booster (11.02 lbs)

Drogue to Main

1026.11 ft·lbf 1252.58 ft·lbf 1182.81 ft·lbf

Main to Landing

14.70 ft·lbf 17.94 ft·lbf 16.94 ft·lbf

Altitude and Drift Predictions

Altitude and drift predictions have been performed for the final vehicle design, and are summarized in Table 23 and Table 24. The values were obtained from the simulation in OpenRocket for the official launch in Huntsville, AL. The vehicle profile is optimized for altitude by adjusting the ballast for various cases of wind speed.

Wind (mph) Ballast Needed (oz) Simulated Apogee Altitude (ft)

0 10 5279

5 9 5274

10 7 5283

15 4 5277

20 0 5289

Wind (mph) Lateral displacement from Launchpad(ft)

0 7

5 560

10 1316

15 1908

20 2682

Verification of System Level Functional Requirements

The verification plan in effect reflects how each requirement to the vehicle satisfies its function. Requirements from the SOW are paraphrased followed by the design feature that satisfies each requirement. Ultimately, each design feature undergoes verification to ensure that it actually meets its requirements. Testing, analysis, and inspection serve as the mode of verification for each feature.

Table 22: Kinetic Energy by Section During Flight

Table 23: Altitude Predictions

Table 24: Drift Predictions


61


Table 25 gives each vehicle requirement, satisfying design feature, and verification method.

Requirement (SOW)

Vehicle Requirement Satisfying Design Feature

Verification Method

1.1 Vehicle shall deliver payload to 5,280 feet AGL

Cessaroni L1720-WT Testing, Analysis

1.2 Vehicle shall carry one official scoring barometric altimeter

Adept A1E, included in the SMD payload

Inspection

1.2.1 Official scoring altimeter shall report the official competition altitude via a series of beeps

Adept A1E Functionality

Testing, Inspection

1.2.2.1

At Launch Readiness Review, a NASA official will mark the altimeter to be used for scoring

Adept A1E can be located easily through clear acrylic body section

Inspection

1.2.2.3

At launch field, all audible electronics except for scoring altimeter shall be capable to turn off

Recovery altimeters can be disabled externally via magnetic arming switch

Testing

1.2.3.1

Official, marked altimeter cannot be damaged; must report an altitude with a series of beeps

Successfully recovery system; sufficient mounting

Testing, Inspection

1.2.3.2

Team must report to NASA official designated to record altitude with official marked altimeter on launch day

This task will be assigned to an appropriate team member

Inspection


62


Requirement (SOW)


Verification Method

1.2.3.3 Altimeter must not report apogee altitude of over 5,600 feet

Cessaroni L1720-WT/ Vehicle Mass

Testing, Analysis

1.3 Launch vehicle remains subsonic from launch until landing

Cesaroni L1720-WT Testing, Analysis

1.4 Vehicle must be recoverable and reusable

Successful recovery system

Testing, Inspection, Analysis

1.5 Launch vehicle shall have a maximum of four independent sections

Vehicle is composed of 3 tethered sections

Inspection

1.6 Launch vehicle shall be prepared for flight at launch site within 2 hours

Launch operations and assembly procedures

Testing, Inspection

1.7

Launch vehicle will remain launch-ready for a minimum of one hour with critical functionality

System runtime capability of at least 2 hours

Testing, Inspection, Analysis

1.8 Vehicle shall be compatible with 8 feet long 1 inch rail (1010) or 1.5 inch (1515)

1010 and 1515 rail buttons attached to vehicle body

Inspection

1.9 Launch vehicle will be launched by a standard 12 volt DC firing system

Cesaroni L1720-WT Igniter

Testing


63


Requirement (SOW)


Verification Method

1.10 Launch vehicle shall require no external circuitry or special equipment to initiate launch

Motor ignition only requires the 12V DC firing system

Inspection

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

Cesaroni L1720-WT Inspection

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

Motor total impulse of 3695.6 N-s

Inspection

1.15

The full scale vehicle, in final flight configuration, must be successfully launched and recovered prior to FRR

Test Launch Schedule Testing

1.15.2 Payload can, but does not have to be, flown during full-scale test flight.

Payload will be flight-ready for the full-scale test flight

Testing

1.15.2.1 If payload is not flown, mass simulators shall be used to simulate payload mass

Payload mass simulator will be available, if needed

Inspection

1.15.2.1.1

Mass simulators shall be located in same location on vehicle as the missing payload mass

Payload mass simulator will be placed in appropriate location, if needed

Inspection

1.15.4 Vehicle shall be flown in fully ballasted configuration during full scale test flight

Nose cone ballast system

Inspection, Testing

1.15.5

Success of full scale demonstration flight shall be documented on flight certification form, by a Level 2 or Level 3 NAR/TRA observer, and documented in

Team mentor (Pat Gordzelik)

Inspection


64


Requirement (SOW)


Verification Method

FRR package

1.16 Maximum amount teams may spend on vehicle and payload is $5000

Projected Budget of ~ $4202

Inspection, Analysis

The vehicle verification statements are listed below. The statement number is followed by the corresponding requirement from the SOW in parenthesis, followed by the statement with the results of the verification.

1. (1.1) With the selected motor (L1720), the vehicle with payload is capable of

reaching 1 mile AGL and has been verified through simulation analysis and test

launches.

2. (1.2) The Adept A1E will serve as the official altimeter and can be verified by

visual inspection in the payload housing.

3. (1.2.1) The Adept A1E has been inspected and tested and to report altitude via a

series of beeps.

4. (1.2.2.1) Inspection verifies that the NASA official can locate and mark the official

altimeter.

5. (1.2.2.3) Testing has verified that the only audible electronics (altimeters in

avionics bays) can be switched off after the vehicle has landed with magnet.

6. (1.2.3.1)Preparation for test launches has verified that the launch vehicle can be

prepared within 2 hours at the launch site

7. (1.2.3.2) Jake, the lead engineer, is delegated the task of reporting to NASA

official for altitude documentation.

8. (1.2.3.3) With the final vehicle mass, the motor is not capable of delivering the

vehicle and payload above 5600 feet. This has been verified through simulation

analysis and testing.

9. (1.3) The maximum velocity achieved by the vehicle is subsonic, and has been

verified through simulated flights and test launches.

10. (1.4) Test launches and analysis along with inspection have verified that the

vehicle is recoverable and reusable when the recovery system functions as

designed.

Table 25: Vehicle Verification Table


65


11. (1.5) visual inspection can confirm that the entire vehicle is comprised of 3

independently tethered sections

12. (1.6) Test launches have verified that the team can prepare the vehicle for flight

within a 2 hour window.

13. (1.7) Power consumption approximations, along with testing at the launch site,

have verified that the vehicle can remain launch ready for a 1 hour minimum.

14. (1.8) Visual inspection verifies that the vehicle is capable of being launched from

either rail options.

15. (1.9) Testing and analysis have demonstrated that the Cesaroni igniters are

capable of being fired from 12VDC firing system

16. (1.10) Inspection verifies that the launch initiation only requires the 12VDC firing

system.

17. (1.11) The Cesaroni L1720 is a commercially available APCP motor as verified

by inspection.

18. (1.12) Inspection verifies the L1720 as meet this impulse requirement.

19. (1.15) Full scale demonstration flight with mass simulator was conducted on

3/14/2013.

20. (1.15.2) Payload was launched on full scale test flight on 3/15/2013.

21. (1.15.2.1.1) Through inspection and analysis, the placement of mass with the

mass simulator accurately reflects that of the SMD payload.

22. (1.15.4) Full scale demonstration flight carried payload mass simulator on

3/14/13.

23. (1.15.5) The full scale demonstration flight has been documented by team

mentor Pat Gordzelik, verifying the vehicle design for flight ready.

24. (1.16) Inspection and cost analysis verify that the maximum amount spent at the

time of FRR is $4007.21, less than the $5000 limit.

Safety and Environment

Safety and Mission Assurance Analysis

Failure Modes and Effects Analysis

Each failure mode for the vehicle has been assigned a Risk Priority Number (RPN). The RPN is a ranking system to determine which failure modes are more likely to occur and require the most attention. The RPN value is calculated by taking the product of three scores: severity, occurrence, and detection. Severity is the importance of the effect upon the project. Severity is scored on a scale from one to five, where one is not severe and five is very severe. Occurrence is the frequency with which a given cause occurs and creates failure modes. Occurrence is scored on a scale from one to five, where one


66


is not likely and five is very likely. Detection is the ability to detect or prevent a given cause. Detection is scored on a scale from one to five, where one is easy to detect and five is not easy to detect.

Launch Vehicle Design Failure Modes

Table 26 provides an updated summary of potential failure modes that could occur during the design of the vehicle, along with associated severity, occurrence, detection, and RPN.

Potential Failure Mode

Causes Failure Effects

Proposed Mitigation

Completed Mitigation

Sev

erity

Occu

rren

ce

Dete

ctio

n

RP

N

Vehicle unstable Lack of

simulations

Unpredictable flight path,

vehicle lands outside of 2500

ft landing radius, rocket

becomes unrecoverable

Run simulations

Completed (11-30-2012)

5 1 1 5

Acrylics does not

withstand forces throughout flight

Ballistic descent

Vehicle unusable, payload unusable

Tensile strength and flight testing

Competed (1-10-2013)

5 1 1 5

Fiberglass does not

withstand forces throughout flight

Weakening of material from unsuccessful

launches

Vehicle unusable

Tensile strength and flight testing

Competed (1-10-2013)

5 1 1 5

Connection between

acrylic payload housing and

upper fiberglass body

tube becomes detached

Friction fitting

Unpredictable flight path, damage to

vehicle body

Research, design, and using shear

pins

Completed (10/27/2012)

5 1 1 5

Fins cause To much drag

Production error,

miscalculation, simulator not

used

Expected apogee

height not obtained

Research, simulations,

and calculations


5 1 1 5


67


Thrust to weight ratio

is less than 5:1

Lack of simulations

Unpredictable flight path,

vehicle lands outside of 2500

ft landing radius, rocket



and calculations


5 1 1 5

Couplers too long or

too short Design error

Early or no separation


and calculations


5 1 1 5

Launch Operations Structure Failure Modes

Table 27 provides an updated summary of potential failure modes that could occur to the structure subsystem during launch operations, along with associated severity, occurrence, detection, and RPN.



Proposed Mitigation


Sev

erity

Occu

rren

ce

Dete

ctio

n

RP

N

Ejection charges

damage air-frame/vehicle components

Premature e-match ignition

Critical systems become

damaged

Calculations, and Proper testing


5 2 1 10

Motor mount fails to

properly retain motor

Lack of epoxy

Damage to internal systems

Structural testing of the motor

mount


5 1 1 5

Rail button failure

Improperly connected to vehicle

body

Unpredictable flight path

Ensure rail buttons are

properly installed and orientated


5 1 1 5

Insufficient component mounting

Improper installation

Potential system

malfunction

Test mounting integrity


5 1 1 5

Airframe stress failure

Lack of testing, ballistic descent

Loss of vehicle functionality, potential loss

of vehicle

Structural testing of airframe


5 1 1 5

Table 26: Potential Failure Modes for the Design of the Vehicle


68




Proposed Mitigation


Sev

erity

Occu

rren

ce

Dete

ctio

n

RP

N

Fin Detachment

Improperly assembled

, or vehicle

experiences ballistic descent

Aerodynamic instability of

vehicle

Ensure that the fins are properly

epoxied


5 1 1 5

Acrylics does not

withstand forces

throughout flight

Ballistic descent

Clear acrylic payload housing section

fractured and shattered in

places ,reusability of

payload section

Size of main parachute

deployment bag (mpdb)

increased, position of mpbd made adjacent to

nose cone


5 2 1 10

Launch Operations Propulsion Failure Modes

Table 28 provides an updated summary of potential failure modes that could occur to the propulsion subsystem during launch operations, along with associated severity, occurrence, detection, and RPN.

Potential Failure

Mode Causes Failure

Effects Proposed Mitigation


Sev

erity

Occu

rren

ce

Dete

ctio

n

RP

N

Igniter does not initiate the

oxidation process for the propellant

Bad Igniter

The vehicle does not launch

Inspect igniter for

concatenation, and always

bring additional igniters for

such an event


5 3 1 15

Propellant‟s oxidation process

does not commence

Bad Igniter

The vehicle does not launch

Use proper igniter, ensue appropriate conditions

when storing propellant


5 1 1 5

Table 27: Potential Structure Failure Modes and RPNs


69




Proposed Mitigation


Sev

erity

Occu

rren

ce

Dete

ctio

n

RP

N

A pressure build-up occurs inside

the motor

Improper motor

assembly, propellant

grain is cracked

Explosion

Inspect the motor, and use

motor assembly checklist


5 1 1 5

Apogee of one mile AGL or greater not achieved

Launch vehicle mass,

launch rod angle

Entire project

Weigh entire launch vehicle

before simulation for launch day,

position launch rod optimally

based on current wind speed and direction


5 5 1 25

Launch Operations Recovery Failure Modes

Table 29 provides an updated summary of potential failure modes that could occur to the recovery subsystem during launch operations, along with associated severity, occurrence, detection, and RPN.

Table 28: Potential Propulsion Failure Modes aand RPNs


70



Causes Failure Effects Proposed Mitigation


Sev

erity

Occu

rren

ce

Dete

ctio

n

RP

N

Drogue parachute

shroud line fails Entanglement

Uncontrollable descent

Research, verify

parachute rating

through testing, and

follow recovery

preparation checklist


3 1 1 3

Main parachute shroud line fails

Entanglement Uncontrollable

descent

Research, verify

parachute rating

through testing, and

follow recovery

preparation checklist


2 1 1 2

No ignition of e-match

Disconnected from altimeter terminal block

Parachute deployment does

not occur

Redundant altimeter system


3 1 1 3

Eyebolt failure Not welded

Altimeter power loss, no main

parachute deployment,

GPS transmitter damage, acrylic payload section shattered, fins

cracking, irreparable body

damage

Verify eye bolt

integrity, and weld


5 1 1 5

Shock cord failure

Shock cord is shredded, or singed from

fire

Untethered vehicle

components, violation of

requirements

Properly fastened

shock cord, verify

rating, and inspect

shock cord


5 1 1 5

Premature black power ignition

Static electricity build up

Premature parachute ejection

Testing, recovery altimeter shielding, and e-match

leads are grounded


5 2 1 10


71





Sev

erity

Occu

rren

ce

Dete

ctio

n

RP

N

Recovery components

damaged

Ballistic descent,

components not shielded

properly, parachutes

develop holes

Failure of recovery system, damage or loss

to vehicle

Use Nomex

cloth and fire

retardant insulation


4 3 1 12

Drogue parachute

comes untied from the quick

link

Unsecured to eye hook with

a quick link

Uncontrolled descent

Secure swivels

along with quick links.


5 1 1 5

Main parachute comes untied from the quick

link

Unsecured to eye hook with

a quick link

Uncontrolled descent

Secure swivels

along with quick links.


5 1 1 5

Drogue parachute

shrouds become entangled

Entanglement with e-match

leads

Uncontrolled descent rate

Cut e-match

leads to fit tube length

loosely


4 2 1 8

Main parachute shrouds become

entangled

Entanglement with e-match

leads


Cut e-match


loosely


5 2 1 10

Uncontrolled separation

Vehicle sections

secured only using friction

fitting

Un-functional recovery system, ballistic descent

Implementing shear

pins


5 1 3 15

PerfectFlite power supply

diminishes

Battery lacking proper

voltage or becomes dislodged

during flight

failure of deployment of parachutes,

Mission failure

Check battery voltage,

and ensure batteries properly

restrained


5 2 1 10

Featherweight Power supply

diminishes

Battery lacking proper

voltage or becomes dislodged

during flight


Mission failure

Check battery voltage,

and ensure batteries properly

restrained


5 2 1 10


72





Sev

erity

Occu

rren

ce

Dete

ctio

n

RP

N

PerfectFlite wired

connections damaged

from using solid core wire, and improper handling


mission failure

Use protected electrical

components, and

braided wire


5 3 1 15

Featherweight wired

connections damaged


Failure of deployment of parachutes,

mission failure

Use protected electrical

components, and

braided wire


5 3 1 15

Drogue parachute

deployment

Black powder charge

undersized, no e-match

ignition

Separation of vehicle

compartments

Over-size powder

charges for main

parachute ejection


3 1 2 6

Main parachute deployment

Black powder charge

undersized, no e-match

ignition

Separation of vehicle

compartments

Over-size powder

charges for main

parachute ejection


5 3 2 30

Rip develops in drogue

parachute

Entanglement of e-match leads with parachute

shroud lines


Cut e-match


loosely


2 1 1 2

Rip develops in main parachute

Entanglement of e-match leads with parachute

shroud lines


Cut e-match


loosely


3 1 1 3

Featherweight not entering flight mode

Improperly sized port

holes

E-match firing, poor to no flight data, parachute deployment, and vehicle damage

Port holes properly

sized


5 1 1 5


73





Sev

erity

Occu

rren

ce

Dete

ctio

n

RP

N

GPS unit damaged

Improperly secured

Loss of tracking signal, difficulty

in recover launch vehicle

New attachment

scheme involving

zip ties and back plate to secure DC40 to drogue

shock cord


3 2 1 6

PerfectFlite not detecting apogee

Improperly sized port

holes

E-match firing, poor to no flight data, parachute deployment, and vehicle damage

Port holes properly

sized


3 1 1 3

Likelihood and Potential Consequences for the Top 10 Failures

Table 30 provides the top ten failure modes for the vehicle based upon the highest RPN value in ascending order, and their associated consequences.

RP

N

Potential Failure Mode Causes Failure Effects

30 Main parachute

deployment

Black powder charge undersized, no e-match

ignition

Separation of vehicle compartments, ballistic descent

25 Apogee of one mile AGL or

greater not achieved Launch vehicle mass,

launch rod angle Mission failure

15 Uncontrolled separation Vehicle sections secured only using friction fitting

Un-functional recovery system, ballistic descent

15 PerfectFlite wired

connections become damaged


failure of deployment of parachutes, mission failure

15 Featherweight wired connections become

damaged


Failure of deployment of parachutes, mission failure

15 Igniter does not initiate the oxidation process for the

propellant Bad Igniter The vehicle does not launch

Table 29: Potential Recovery Failure Modes and RPNs


74


RP

N


10 Vehicle unstable Lack of simulations ran Unpredictable flight path, vehicle lands outside of 2500 ft landing radius, rocket


10 Ejection charges damage

air-frame/vehicle components

Premature e-match ignition Critical systems become damaged

10 Main parachute shroud lines become entangled

Entanglement with e-match leads


10 PerfectFlites power supply

diminishes

Battery lacking proper voltage or becomes

dislodged during flight

Failure of deployment of parachutes, Mission failure

List of Personnel Hazards

Table 31 provides an updated 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 Consequence Mitigation Action

Laceration

Knives, routers, saws,

file, Dremel tool

Medium Serious injury or

death

Follow safety protocols, proper

tool and equipment use, personal safety attire, refer to

operators manual

Discontinue all operations,

apply first aid, contact EMS

Burns

Chemicals (FFFFg,

fiberglass resin),

welders, soldering Iron

Medium Minor to serious

injury



operators manual



Respiratory Damage

Chemicals (epoxy,

solder), fumes, fiberglass

Low Brain damage or

death


tool and equipment use, personal safety attire, consult

MSDS



Table 30: Top 10 Failure Modes by RPN


75



Vision Damage

Welders, fiberglass, grinders, projectile

debris

Low Partial to complete blindness

Use of goggles, force shields,

consult MSDS, first aid kit

available, refer to operators manual


apply first aid, contact EMS, use eyewash

Allergic Reaction Epoxy,

chemicals, fiberglass

Low

Loss of respiration,

inflammation (Internal & External)

Use of gloves, consult MSDS,

first aid kit available


apply first aid, contact EMS,

administer antihistamine, safety shower

Hearing Damage

FFFFg, Grinders, Ignition, Routers

Low Partial to complete deafness

Ear muffs, consult MSDS,

first aid kit available, refer to operators manual



Dismemberment

Projectiles, Saws,

Launches Low

Permanent injury or death

Make sure proper safety measures

are taken, operators manual


apply first aid, and contact

EMS, tourniquet

The material safety data sheet (MSDS) that the manufacturer provides contains information about the material in consideration. It is comprised of 16 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. According to the safety plan, a binder containing all the MSDSs is always made available for personnel and brought to every launch. 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 component used during the project are kept in an operator manual 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.

Table 31: Potential Hazards to Personnel


76


The team is aware that the FAA must be notified of planned launch activities. For educational outreach events, notification to the closest airport within five 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. The team has obtained flight waivers for full scale launches. The flight waiver went into effect on December 15, 2012, and it lasts for one year. The conditions allow for CFR 101.25 (e) and 101.25 (f,g), which is located in Appendix D to be waived. Flight operations occur between sunrise and sunset within a controlled airspace from the surface to 5,280 feet above ground level, 6,569 feet mean sea level (MSL). Operations are coordinated prior to launch dates. The launch location is within a one nautical mile radius of 32.212954N/-0.98.092861W, on the Hunewell Ranch in Stephenville, Texas. To ensure that the conditions from the flight waivers are followed, a procedural checklist has been devised and implemented along with a pre-mission launch briefing which occurs prior to every launch. The flight waivers are located inside the launch procedures binder, which is brought to every launch. (Reference in Appendix E) 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 vehicle 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) 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) Launch Procedures and Checklists A summary of legal risks that could occur during the course of the project appears in Table 32.

Risk Likelihood Severity Consequence Mitigation

FAA Violations Low High Legal

Repercussions Adhering to Regulations

NAR/TRA Violations

Low High Legal


Damage of Property

Low High Legal

Repercussions Insurance


77


OSHA Violations

Low High Legal


Personal Injury Low High Legal

Repercussions

Redundant Calculations and

Safety Preparedness

Environmental Concerns

In the event of an unrecoverable or damaged vehicle, certain materials could be left exposed to the environment. The biodegradability of each material used effects the impact on the surrounding ecosystem. Much of the information concerning the hazards posed to the environment and ecology is available in the individual MSDSs. The effects of materials used in the construction and launch of the vehicle are summarized in the following Table 33.

Material Prevalence Mode of Biodegradability Impact on Environment

Ammonium Perchlorate

Motor propellant Highly water soluble Iodization of local water

table

Black Powder Ejection charges Remains solid No known impact

Epoxy 2M DP420 Fiberglass

connections, sealed couplings

Decomposition begins within fifteen months

Leaching to local water table

Oil-Based Spray Paint:

External fiberglass structural components

Soluble Leaching to local water

table

Clear Acrylic: Payload bay Soluble Leaching to local water

table

Fiberglass Structural

components Remains solid

No known environmental impact; may

pose ecological hazard

Kevlar Deployment bag Remains solid No known

environmental impact; may pose ecological hazard

Aluminum Motor tube, rivets,

battery casing, Highly water soluble

Long-term degradation products

Cellulose Recovery wadding; Remains solid Long-term degradation

products

Steel Attachment hardware,

ballast system Remains solid

Leaching to local water table

Copper: Avionics bay lining, e-

match lead wires Highly reactive in air or

moisture Long-term degradation

products

Sulfuric Acid Batteries Highly water soluble Long-term degradation

products; may pose ecological hazard

Kevlar Shock harnesses Remains solid No known impact

Table 32: Legal Risks


78


Material Prevalence Mode of Biodegradability Impact on Environment

Silicon Parachutes, batteries Reactive in air or moisture Irritating vapors form

Rip-Stop Nylon: Parachutes, shear

pins Remains solid No known impact

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 weather, flight waivers have been obtained so that alternate test dates are easily rescheduled. Launch dates can be viewed in the testing timeline in Figure 89. Environmental factors such as surrounding flora, fauna, or sedimentary projections could cause the launch vehicle to become unrecoverable. These risks are outlined in Table 34.

Risk Likelihood Severity Consequence Mitigation

Poor weather High High Delay in testing Multiple test

dates,

Environment prevents recovery

Medium Medium Possible loss of

Vehicle

Survey launch site, recovery

tools

Burn ban in effect

Low High Delay in testing Multiple test

locations

Payload Integration The payload is designed to integrate into the acrylic payload housing structure of the launch vehicle with ease. The payload is constructed on a payload framework which consists of a forward payload bulkhead, a rear payload bulkhead, and a rectangular aluminum frame.

Integration of the Payload with the Launch Vehicle The payload is designed to integrate into the acrylic payload housing structure of the launch vehicle with ease. The payload is constructed on a payload framework which consists of a forward payload bulkhead, a rear payload bulkhead, and a rectangular aluminum frame. The rear payload bulkhead is friction fit into the rear coupler (booster to payload housing) which houses the drogue avionics bay. A rectangular frame constructed from

Table 33: Effects of Materials used in Construction and Launch

Table 34: Environmental Factors Hendering Recovery


79


four aluminum angle rails will also attach to the rear payload bulkhead (Figure 51) with two screws.

The forward payload bulkhead is friction fit into the forward coupler (payload housing to upper body airframe) which houses the main parachute avionics bay. This forward payload bulkhead has a slot, created by a bracket (as seen in Figure 52), where the open end of the aluminum frame will seat upon payload is installation.

Figure 51: Rear Payload Bulkhead

Figure 52: Bracket on Forward Payload Bulkhead


80


The payload is integrated with the launch vehicle by securing the payload framework to the rear payload bulkhead with two screws as seen in Figure 53.

Next, the rear coupler is inserted in the lower section of the acrylic payload housing structure, ensuring the opposite end of the payload framework seats properly into the forward payload bulkhead‟s framework slot as shown in Figure 54.

Figure 53: Securing Payload Framework to Rear Bulkhead

Figure 54: Framework Fitting into the Bracket


81


The coupler secures with six screws; from the exterior of the payload housing, through the coupler, and into the rear payload bulkhead. At this point, the physical integration of payload is complete as depicted in Figure 55.

Step Procedure Action taken 1 Secure Payload Framework

to Rear Payload Bulkhead Two screws installed

2 Aligning aluminum Framework with static bulkhead slot

Rotate Payload Framework Until Aluminum Frame is Aligned

3 Seating Aluminum Framework into static bulkhead slot

Press Payload Framework into Forward Payload Bulkhead Slot

4 Securing Payload Framework

Install Six Screws From the Exterior of the Vehicle into the Rear Payload Bulkhead

Compatibility of Elements

The payload is designed to operate within the launch vehicle, thus the payload must be compatible with the launch vehicle. To ensure compatibility, the payload is designed to meet the specifications of the internal payload housing structure. All payload circuits mounted on the framework are oriented to provide sufficient clearance between the payload housing structure‟s internal diameter and the individual components. Portholes are installed to the payload housing structure to allow ambient atmospheric pressures to equalize within the payload housing structure to allow accurate readings of the BMP180 pressure sensors. Clear acrylic was chosen as the material for the payload housing structure. This allows for use of the camera, solar irradiance, and UV radiation sensors from within the launch vehicle. The acrylic is UV-T specification to allow approximately 85% of UV light to pass through the material without being filtered.

Figure 55: Fully Assembled SMD Payload Section

Table 35: Procedures for Installing Payload


82


Payload Interface Dimensions The payload framework consists of six components; two bulkheads (rear payload

bulkhead and forward payload bulkhead) and four aluminum angle rails. The rear payload bulkhead is a cylindrical piece of PVC which is 1.0 inches in height with a radius of 2.55 inches. This will provide a precision fit with the interior of the rear coupler which will be epoxied into place. The static bulkhead is identical with the exception of a 0.55 inch wide by 0.55 inch deep slot oriented on the center of the bulkhead and continuing across the width of the bulkhead surface. This will provide a 0.1 inch tolerance between the slot and aluminum framework; ensuring a precision fit, while allowing easy installation. The aluminum frame is constructed from four pieces of aluminum angle. The aluminum angle is 0.5 inches wide by 0.5 inches height by 0.125 inches thick. Once constructed, the aluminum frame is a rectangular box 29.0 inches long by 4.85 inches wide. The angles are joined with aluminum welds.

Item Dimensions Bulkheads 1.0 inches Height

2.55 inches Radius

Static Bulkhead Slot 0.5 inches Height 0.75 inches Depth 4.875 inches Width

Aluminum Angle 0.5 inches Height 0.75 inches Width 0.125 inches Thick

Long Rails 20.5 inches

Short Rails 4.875 inches

Rail Separation Distance (Inner)

3.875 inches

Rail Separation Distance (Outer)

4.875 inches

Payload-Housing Integrity

The payload-housing for the SMD payload is the payload housing structures acrylic body tube. The acrylic was subjected to a crush test and survived 1,500 foot pounds of force with no failures. The payload housing has also been test launched in at least six subscale launches. To preserve the integrity of the payload housing two specific actions are required.

Table 36: Payload Interface Dimensions


83


One action is the use of nylon bushings on each screw securing the payload housing to the couplers; to prevent the acrylic from shattering upon landing. The nylon bushings are more ductile than both the acrylic and metal screws and dissipate the forces placed on the screws during operation of the vehicle. The other action is ensuring payload housing sits flush against the coupler collars to prevent applied forces from being unevenly distributed. Also of note, the acrylic used to house the payload is of UV-T specification; allowing approximately 85% of UV light to pass through the material without being filtered. This allows the use of the UV sensors internally.

Integration Demonstrated

The payload is integrated with the launch vehicle by securing the payload framework to the rear payload bulkhead with two screws. First, the payload framework is secured to the rear payload bulkhead. This bulkhead is already friction fitted within the rear coupler as seen below in Figure 56.

Once the payload framework is secured to the rear payload bulkhead, the assembly should look like Figure 57.

Figure 56: Payload Framework Secured to Rear Payload Bulkhead


84


The payload framework and rear coupler are then inserted within the payload housing structure and positioned to allow the payload framework to seat into the slotted bracket on the forward payload bulkhead as shown in Figures 58 and 59.

Figure 57: Payload Framework Secured

Figure 58: Payload Integration

Figure 59: Payload Integration


85


The payload is now integratged, and needs only to be secured with six screws; from the exterior of the payload housing, through the coupler, and into the rear payload bulkhead as seen in Figure 60.

Figure 60: Integrated Payload


86

IV) Payload Criteria


Experiment Concept The payload design fulfills the requirements of the SMD payload. It records

measurements of pressure, temperature, relative humidity, solar irradiance and

ultraviolet radiation; these measurements are stored onboard and transmitted to the

team‟s ground station. During the flight and post flight the data is analyzed in order to

determine correlations between the different measurements.

The payload housing is made of clear acrylic. The sensors take their readings from

within the housing, eliminating the need to eject. This design emphasizes safety in

terms of the electronics and ground observers. A clear acrylic housing does have some

disadvantages in terms of the payload experiment; the housing partially filters UV

radiation. This filtering is discussed in more detail in the Precision of Instrumentation

section of this document.

The payload includes a camera leveling system comprised of two servo motors, an

accelerometer (tilt sensor), and a photographic camera. This autonomous real-time

camera orientation system (ARTCOS) is responsible for meeting requirement 3.1.3.6 of

the SOW. The major tests of the ARTCOS are for structural stability of the component

mounting, speed of the servos, and the reaction time of the accelerometer.

Science Value

Payload Objectives

The main objectives of the payload are to record and store atmospheric and GPS data, transmit these measurements to a ground station, and capture photos during the rocket's flight. The NASA USLI SLP states additional objectives which are addressed in the following section.

Payload Success Criteria

The payload experiment is deemed successful if it obtains and transmits valid atmospheric data and flight imagery. Validity of the measurements is based not on their accuracy, but rather on whether they lie within the established confidence intervals for their particular sensors. Should any data lie outside this acceptable range, it can be considered either a mechanical error of the sensor or a computational error within the software which renders data invalid.


87


Validity of the imagery is based on whether the orientation of the image meets the criterion of proper orientation, with the ground of the scenery at the bottom of the frame and the sky at the top of the frame. Successful operation of the ARTCOS system will determine whether captured images meet this standard. If the payload does not record at least two photos during the rocket's descent and three photos after the rocket's landing the imagery system is deemed unsuccessful. Validity of the data transmission is based on whether the information received at the ground station corresponds with the recorded measurements. Should any values differ, it can be assumed that bits were lost during the transmission and the values received are consequently invalid. These transmissions should occur at a specified frequency of 200 megahertz. If the ground station does not receive information every five seconds, the communication system is deemed unsuccessful. Table 37 summarizes the payload objectives, their science value and success criteria.

Payload Objective Science Value Success Criteria

Gather Atmospheric and GPS Data

Representing Change in Atmospheric Variables Depending on Altitude

Collected Valid Data

Store and Transmit Atmospheric and GPS Data

Modeling Rocket Flight

Collected at Required Frequency, Transmitted Throughout Flight

Camera Orientation Test Multi-Servo Orientation Device

Two Pictures During Descent and Three after Landing, Correct Orientation

Experimental Logic, Approach, and Method of Investigation The SMD payload gathers data from approximately 5,280 feet above ground level down to the landing site. Data gathered includes 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 variables. These two calculations aid in the development of 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.

Table 37: Payload Objectives


88


Regression Model

With a large number of samples ranging across the various test and demonstration flights, the data plots can be analyzed statistically to determine a model that accurately represents relationships between the different variables. Using R-based software, a regression model of the form Y = Xβ + ε is computed. In these models, a chosen response variable (Y) is modeled against a column matrix of the other variables (X) multiplied by their derived coefficients (β) and a random error term (ε). Assuming that the response variables and error terms meet requirements of constancy and normality in their variance, these models can be tested for their validity. A t-test will suggest if any column in the variable matrix X should be removed. A stepwise model can suggest the optimal subset of variables to be included in the model. Each model that produced the R2 statistic will show how well the predictive values of the created model matches with the experimental data. Furthermore, the residual sum of squares of the error terms demonstrates how far the model deviates from the actual data. These tests, among others, can be used to create, refine, and validate any models.

Correlations between SMD Sensor Readings The distributions of each variable aid in finding covariance in pairs of variables; therefore, the correlation between those variables can subsequently be attained. After determining the correlation, the model can evolve as necessary to provide more accuracy. Using R-based software, any data set can be analyzed by individual variables. Variance matrices and standard deviations from each variable can be combined in well-defined formulae to find covariance and correlation

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


89


Test and Measurement, Variables, and Controls

Measurements will be taken by the payload sensor circuitry. Although some measurements might be taken at the ground station, these will be used for predicting the rocket's trajectory. The only data considered in the formal analysis will come from the onboard payload circuitry. The variables included in the measurement process are altitude, temperature, atmospheric pressure, relative humidity, solar irradiance, UV radiation, and random electrical noise in the payload circuitry. 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 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. Using these models, a select few measurements made on the ground level could accurately predict weather conditions at the launch site. This would enable amateur rocket enthusiasts to more precisely predict the exact flight path of a launch vehicle. 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 A detailed preliminary experiment has been conducted to find any initial models or relationships against which to test. Using the same processes mentioned above, over five gigabytes of atmospheric data from NOAA was analyzed. This data was limited to a subset corresponding to Stephenville, Texas to eliminate any confounding variables relating to geographical differences. The results are as follows: CC = 478.5 – 1.028 ILW – 0.3985 PW – 0.003413 P – 614.1 SH ILW = 330.3 – 0.5912 CC – 0.2285 PW – 0.002283 P PW = -402.4 – 0.243 CC – 0.2496 ILW – 0.01109 ISW + 0.004428 P – 0.0595 RH + 1349 SH RH = 0.5537 PW + 12910 SH


90


where “CC” denotes cloud coverage, “ILW” denotes long-wave solar irradiation, “ISW” denotes short-wave solar irradiation, “P” denotes atmospheric pressure, “PW” denotes precipitate water, “RH” denotes relative humidity, and “SH” denotes specific humidity. After establishing these models, a subset of the NOAA data corresponding to Huntsville, Alabama was used to validate the model. These calculations, among others, suggested these models to be valid. Concerning accuracy, atmospheric variables outside the payload's domain of measurement were considered. These extraneous variables were precipitate water, cloud coverage, ground roughness, specific humidity, geopotential height, and wind velocity measurements. As expected, many of these variables had nominal effects on the variables considered by the payload. Although the variables of cloud coverage and precipitate water demonstrated significant impact on the variables measured by the payload, their values can be either substituted or accurately calculated using versions of the formulae listed above.

Payload Design The payload consists of two main systems; the Atmospherics Data Gathering System

(ADGS), the Autonomous Real-Time Camera Orientation System (ARTCOS). The

payload records measurements of pressure, temperature, relative humidity, solar

irradiance and ultraviolet radiation; these measurements are stored onboard to a micro

SD card. The payload does not eject from the vehicle, but rather takes all readings

internally through a clear acrylic housing. The electronics are mounted on an aluminum

rail framework. The camera is kept level throughout the descent phase of the flight by

the ARTCOS, which includes a dual servo motor mechanism. Camera images are

stored to a micro SD card. A 900 megahertz XBee S3B transmitter allows a 28 mile

range of wireless transmission. A custom Printed Circuit Board (PCB) minimizes the

space utilized by the electronics and improves the signal integrity between the

components. Advanced Circuits in Aurora, Colorado has sponsored the PCB

manufacturing. A video camera records flight footage for educational outreach

purposes.

A CAD rendering of the payload design is shown in Figure 61.


91


The payload is built and was flight tested in the FRR full scale launch. The constructed

payload is shown in Figure 62.

Figure 61: Payload Design


92


The payload electronics are soldered on to PCBs which are made of FR-4 which is a

woven glass fabric with an epoxy resin system. Precise characteristics of FR-4 are

shown in Table 38.

Figure 62: Constructed Payload


93


The PCBs are mounted to 1/16 inch thick vertical aluminum rails through the use of

gauge four screws at each of the corners of the PCBs, and corresponding gauge four

nuts secure the screws. The rails are welded at the top and bottom for structural

stability. The bottom is screwed on to a bulkhead which is a cylindrical piece of PVC 1.0

inches in height and a radius of 2.55 inches. The rails fit securely into a slot on the

second bulkhead. Payload integration is described in further detail in the Payload

Integration section of this document.

ADGS

The ADGS consists of two PCBs and a box framework for the UV and solar irradiance

sensors. The front PCB contains an Arduino Mega 2560 microcontroller, a TAOS

TSL2561 lux sensor, a Locosys LS20031 GPS module, a Bosch BMP180

pressure/temperature sensor, an XBee XSC S3B transmitter, and an Adafruit micro SD

card reader/writer. The PCB schematic, containing the exact dimensions of the board

Table 38: FR-4 Characteristics


94


and the placement of the components, is shown in Figure 63. The green lines represent

the board itself, the red lines represent traces on the top of the board, the blue lines

represent traces on the bottom, and the yellow dots represent vias.

Figure 63: ADGS PCB Schematic


95


The components are screwed into place with gauge two screws for structural stability.

An actual photo of the board is shown in Figure 64. The picture does not include the

LCD screen.

The second PCB included the ADGS holds two nine-volt batteries, three relays, and a

magnetic arming switch. The board is responsible for providing power to the ADGS and

acts as a switching circuit for the entire payload. This PCB is located on the backside of

the other PCB and uses the same screws for mounting. The board has an L-shape

Figure 64: ADGS PCB


96


because the Arduino Mega is located on the back of the other board. Rubber gromits

provide separation between the two boards. The PCB schematic, including the exact

dimension and component placement, is shown in Figure 65.

The batteries are placed in plastic mounts which are screwed down and secured with

metal zip-ties. An actual picture of the power board with components mounted is shown

in Figure 66.

Figure 65: Power PCB Schematic


97


The box framework for the solar irradiance and UV sensors is constructed of fiberglass

and screwed into the aluminum rails. The fiberglass sheets are connected with epoxy.

Each of the sensors is screwed onto the fiberglass and wired to the main ADGS PCB. A

CAD rendering of the box framework with exact dimensions is shown in Figure 67.

Figure 66: Power Board


98


Measurements from the ADGS are transmitted to the team ground station. The ground

station includes a handheld Yagi antenna connected to a 900 MHz XBee XSC S3B

which is connected to a laptop running a Java application. The Java application saves

and displays the readings received by XBee. The CDR design used a MATLAB GUI

instead of a Java GUI. The change has been made due to the extensive multi-threading

capabilities of the Java JDK. The ground station software is still in development, the

current GUI is shown in Figure 68. Real time graphs of the sensor readings are

displayed.

Figure 67: Solar Sensor Array

Figure 68: Real Time Sensor Readings


99


ARTCOS

The entire ARTCOS includes two PCBs and the dual servo mechanism. The front PCB

contains two Arduino Pro Mini 328 microcontrollers, a Honeywell HIH4030 humidity

sensor, an Adafruit micro SD card reader/writer, and a 3.3 Volt voltage regulator. This

board is responsible for controlling the orientation of the servos, saving the pictures, and

sending the humidity data to the ADGS. The PCB schematic with exact dimensions and

component placement is shown in Figure 69.

The ARTCOS PCB is incomplete. There was an error in shipping and the board will not

arrive until after this document is submitted; therefore, there is no picture of the actual

PCB with placed components. The board will be shipped and completed by the FRR

presentation.

A second PCB holds a nine-volt battery for the servos, a nine-volt battery for the

ARTCOS control board, and a five-volt voltage regulator. This board is placed behind

the other ARTCOS PCB and mounts to the same screws. Rubber grommets separate

the two boards. This board connects to the ADGS power board so that the payload can

be activated from a single switch. The PCB schematic is shown in Figure 70. The

dimensions are the same as the previous board. The batteries are secured in the same

fashion as the ADGS power board (except the battery holders hang over the board

slightly).

Figure 69: ARTCOS Control PCB Schematic


100


In order to control the orientation of the camera throughout the descent of the flight, the

camera is attached to two servos whose angle is determined by readings from an

accelerometer mounted on top of the camera. Custom designed and built PVC

mounting brackets hold all the pieces together. A fiberglass mount secures the entire

piece to the aluminum rail. This mount is a major improvement from the previous design

and was deemed necessary from previous flight failures. A CAD rendering of each of

the individual pieces is shown in Figure 71.

Figure 70: ARTCOS Powerboard PCB Schematic


101


An actual picture of the entire ARTCOS mechanism mounted to the aluminum rail is

shown in Figure 72. Wire entanglement is a major issue to consider for functionality of

the ARTCOS. The wires are managed such that they run on the inside of the aluminum

rails.

Figure 71: ARTCOS Explosion Assembly


102


The exact wiring is found in the PCB schematic. A visual representation of the

connections between components for the overall payload is shown in Figure 73.

Figure 72: ARTCOS Assembly


103


A list of the electrical components onboard the payload is compiled in Table 39.

Purpose Breakout Board Distributor

Part Number Interface Dimensions (L x W x H)

Input Voltage

Current Draw

Flight Computer

Sparkfun Arduino 2560-R3

N/A 2.125 x 4.3125” 7 – 12V 20 - 200mA

Flight Computer

Sparkfun Arduino Pro Mini 328

N/A 0.7 x 1.3” 5 – 12V 150mA

Humidity Sensor

Sparkfun HIH4030 Analog 0.75 x 0.30” 4 – 5.8V 200μA

Temperature Sensor

DSS Circuits

BMP180 I2C 0.625 x 0.5” 1.8 - 3.6V 3 – 32μA

Pressure Sensor

DSS Circuits

BMP180 I2C 0.625 x 0.5” 1.8 - 3.6V 3 – 32μA

Figure 73: Component Connections


104


UV Sensor Apogee SU-100 Analog 0.925 x 0.925 x 1.08”

0V 0A

Lux Sensor Adafruit TSL2561 I2C 0.75 x 0.75” 2.7 – 3.6V 0.5mA

Pyranometer Apogee SP-110 Analog 0.925 x 0.925 x 1.11”

0V 0A

Wireless Transmitter / Receiver

Digi XBee-PRO XSC S3B

Serial 1 x 1” 2.4 – 3.6V 215mA

XBee Spacer Parallax 32403 N/A 1.16 x 1.0 x 0.58”

0V 0A

GPS Locosys LS20031 Serial 1.18 x 1.18” 3 – 4.2V 29mA

Camera Adafruit VC0706 Serial 1.26 x 1.26” 5V 75mA

Servo Servo City HS55BB+ Mighty Micro

Pulse Width Modulation

1.14" x 0.51" x 1.18"

5V 240mA

Tilt Sensor Sparkfun ADXL345 I2C 1 x 0.5” 3.3V 40μA

Micro SD card

Wal-Mart 3FMUSD16FB-R

N/A 2.2 x 0.3 x 3.4” N/A N/A

Micro-SD PCB

Adafruit 254 SPI 1.25 x 1 x 0.15” 3 – 5V 150mA

LCD Sparkfun 11062 SPI 1.5 x 2.5” 3.3 – 6V 108 – 324mA

Voltage Regulator

Mouser LD1085V50 N/A 0.25 x 0.25” 5 – 30V -

Battery Ultralife U9VLBP N/A 1.81x1.04x0.69” 5.4 – 9.9V Max 120 mA

Precision of Instrumentation and Repeatability of Measurements

Ultraviolet Radiation

The SU-100 datasheet specifies the range of the light spectrum that the sensor is sensitive to. The sensor is most responsive to UVA (320 – 400 nanometers) and UVB (280 – 320 nanometers). The SU-100 datasheet provides the sensor‟s spectral response in Figure 74. The datasheet lists the absolute accuracy at 10% and the repeatability at 1%.

Table 39: Payload Electrical Components


105


The SU-100‟s accuracy is not the only factor affecting the accuracy of the UV measurements obtained by the payload. The UVT clear acrylic housing filters some of the UV rays. The data sheet for the UVT acrylic specifies the spectrum and the percentage of UV that is filtered by the UV. The filtering is specified in Figure 75.

Solar Irradiance

Figure 74: SU-100 Spectral Response

Figure 75: Acrylic Sheet UV Transmission Curves


106


The SP-110 measures solar irradiance. The sensor is responsive to the range of the light spectrum from 300 nanometers to 1,100 nanometers, and it is most sensitive at approximately 975 nanometers. The SP-110 datasheet shows the sensor‟s spectral response in Figure 76. The datasheet lists the absolute accuracy to be 5% and the repeatability to be 1%.

The accuracy of each of the payload sensors, according to their datasheets, is listed in Table 40.

Purpose Product Precision Barometric Pressure BMP180 ±0.017psi

Temperature BMP180 ±1.8° F

Humidity HIH4030 ±3.6% RH

Solar Irradiance SP-110 ±5%

Solar Irradiance TSL2561 ±5%

Ultraviolent Radiation SU-100 ±10%

GPS LS20031 ±9.84ft

Accelerometer ADXL345 ±4.3mg

Official Altimeter Adept A1E ±1ft

Figure 76: Spectral Response for Apogee Pyranometer

Table 40: Payload Sensor Precision


107


Flight Performance Predictions

The payload was flown in the FRR full scale launch. The results of the launch show

several failures and several successes. Solutions to the failures have been conceived

and will be implemented before the final launch in Huntsville. The successful results of

the full-scale launch give accurate flight performance predictions. These predictions are

specified in this section.

Atmospheric data is saved to a micro SD card onboard the payload and transmitted to

the ground station over a 900 MHz XBee XSC S3B radio. Throughout the flight,

telemetry packets containing the sensor data transmitted every 1.2 seconds. Pressure

measurements are gathered with a BOSCH BMP180. The payload recorded pressure

throughout the flight. The pressure dropped at launch and rose steadily after apogee. A

graph of the recorded pressure is shown in Figure 77.

Pressure data was recorded in the test launch conducted prior to the CDR. The data

from the most recent launch produces a smoother, more accurate graph. The data from

the CDR launch is shown in Figure 78.

20

21

22

23

24

25

26

27

28

29

30

Pressure (inHg)

Figure 77: Pressure Data


108


Temperature was also recorded throughout the flight. The temperature was already high

at liftoff. Upon further inspection, it is clear that the temperature recorded by the sensor

steadily rose from the time the rocket was placed on the launch pad. The temperature

data from the time the rocket was placed on the launch pad is shown in Figure 79.

25.5

26

26.5

27

27.5

28

28.5

29

CDR Pressure (inHg)

86

87

88

89

90

91

92

93

94

Launch Pad Temperature (oF)

Figure 78: CDR Pressure Data

Figure 79: Launch Pad Temperature


109


The same phenomenon occurred in the test launch prior to the CDR. The root of this

problem was determined to be the location of the sensor in relation to the sun. The

sensor was then placed on the opposite side of the PCB to shield it from the light. Since

the new placement of the sensor did not solve the problem, the new hypothesis is that

inside the clear acrylic the temperature rises and the effects of altitude on the sensor

are minimal. An external sensor would be a more appropriate location for a temperature

sensing device. The Recorded temperature for the flight is shown in Figure 80. The

temperature increases upon launch and does not decrease after apogee.

Humidity is measured by a Honeywell HIH4030 sensor. Humidity data for the flight was

not accurate. A bad connection is suspected to be the culprit for the bad humidity data

recorded during the flight. The humidity sensor is located on the ARTCOS Control PCB,

which is the PCB that has not been shipped on time. The circuit flown in the full-scale

flight was a prototyping circuit and the connections were not optimal. However the

HIH4030 humidity sensor was working in the launch prior to the CDR; that data is

shown in Figure 81.

91

91.2

91.4

91.6

91.8

92

92.2

92.4

92.6

92.8

93

Temperature (oF)

Figure 80: Temperature Data


110


Solar irradiance is measured by a TAOS TSL2561 lux sensor and four Apogee SP110

Pyranometers. The lux sensor is placed on the lower PCB and the measurements

gathered by the sensor are adversely affected by the orientation of the payload, while

the Pyranometers are placed facing every 45 degrees so that the orientation of the

payload in relation to the sun does not affect the accuracy of the measurements

gathered by the sensors. The readings from the lux sensor are shown in Figure 82.

0

5

10

15

20

25

30

35

40

45

50

2:38:24 PM 2:52:48 PM 3:07:12 PM 3:21:36 PM 3:36:00 PM 3:50:24 PM 4:04:48 PM 4:19:12 PM 4:33:36 PM

CDR Launch Pad Humidity (%RH)

-500

0

500

1000

1500

2000

2500

3000

Lux

Figure 81: CDR Launch Pad Humidity

Figure 82: Lux Sensor Data


111


The readings from the four pyranometers are shown in Figure 83. When taking the

maximum of the four sensors for each time step, the graph appears to show a

correlation between altitude and solar irradiance.

Solar irradiance measurements were taken in the launch prior to the CDR. That test

flight only contained two SU110 placed facing 180 degrees from each other. The data

from the CDR launch is shown in Figure 84.

0

100

200

300

400

500

600

700

800

Series1

Series2

Series3

Series4

Solar Irradiance (W/m2)

Figure 83: Solar Irradiance


112


The UV is measured by four Apogee Instruments SU100 sensors. The payload did not

record any UV measurements throughout the flight. The payload did not record UV data

for the CDR flight either. The problem then was that the acrylic was the UV blocking

type and new clear acrylic was ordered. The new UVB acrylic was lab tested and

showed a minimal amount of UV blocking, between 10 and 15 percent. The hypothesis

for the UV sensor failure for this flight is that the sun was low in the sky and the small

amount of UV light hitting the payload was blocked by the acrylic. While this is the initial

hypothesis, more testing needs to be performed before a definite reason for the failure

can be decided.

A Locosys LS20031 GPS module records the latitude, longitude, and MSL altitude. The

GPS data was recorded throughout the flight. A graph of the GPS flight data is shown in

Figure 85. The altitude has been scaled down for graph readability.

-100

0

100

200

300

400

500

600

700

800

900

Figure 84: CDR Solar Irradiance Data

CDR Solar Irradiance (W/m2)


113


GPS data was recorded during the launch prior to the CDR. The data from that flight is

shown in Figure 86.

Figure 85: GPS Flight Data


114


An ARTCOS failure was suspected preflight as the servos were not moving when the

payload was activated. The camera was taking pictures and storing them to a micro SD

card. Post flight, a disconnected wire was found to be the cause of the servo failure.

Pictures were stored throughout the flight and after landing, these pictures are shown in

Figure 87. The image resolution was set low in order to speed up the frequency of

picture taking. In future flights the image resolution will be increased.

Figure 86: CDR GPS Flight Data


115


Figure 87: ARTCOS Flight and Landing Photos

The Keyfob video camera‟s battery died and video of the flight was not recorded. The

hypothesis is that the batteries were not charged fully during pre-launch preparations.

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


116


altimeter is the GP-23A Alkaline Lighter Battery. The official altimeter did not record an

official altitude. The hypothesis is that the altimeter is broken, or the port holes for the

payload were incorrectly sized. The payload has five quarter inch port holes, located

eight inches from the upper body airframe. Another hypothesis is that the orientation of

the sensor was incorrect. The sensor is mounted to the inside of the aluminum rail, just

above the cube of solar sensors. Upon landing the altimeter continually beeped an

altitude of 48 feet.

Approach to Workmanship

Construction of the payload requires coordination of several sub-teams. The structural team designs a framework to ensure compatibility between the payload and the launch vehicle. The Electronics Hardware Team designs a power system to power the payload and verifies the physical wiring and interfacing of the components. This involves the planning and designing of wiring schematics for connecting the various components of the payload as well as the power management system to ensure that all components have adequate power and proper voltages to operate. Additionally, the Electronics Hardware Team designs the board layouts and solders the electrical connections between components. The Electrical Software Team programs the payload and ensures functionality of the sensors and components. This includes compiling function libraries for all the sensors to ensure functionality. It also includes creating code to allow them all to be used as a system. Interface software includes the use of I2C, SPI, and USART.

Test and Verification Program

All payload component functionality is analyzed base on 3 areas: hardware, software,

and data analysis.

The testing and verification of requirements has been performed in stages. The first

stage incorporated testing and verifying the wiring of the components on a breadboard.

Then perfboard testing began. At the end of perfboard testing, prototype one was

complete and was flight tested. The results of the prototype one test flight are in the

CDR. Once verification of the perfboard wiring was complete, design, testing, and

verification of the PCBs began. First, the PCBs were designed on the computer using

PCB artist. The designs were then sent to Advanced Circuits, who is sponsoring all of

the team‟s PCB manufacturing. The boards were printed and shipped back. Once

received, parts were soldered onto the boards.

The first stage of PCB development includes breakout boards, which are prefabricated

circuits for each of the individual sensors. This is the stage of development the team is

at currently. The payload is functional with the breakout boards connected to the custom

designed PCBs with the exception of the ARTCOS control PCB which was delayed in


117


shipping. While the current design fulfills all the requirements and is flight ready, there

are still improvements that can be made to the system. The next phase of PCB

development includes moving away from the breakout boards and using surface mount

technology to improve the signal integrity between components and decrease the size

requirements of the boards. These new boards will be used in the final design; however,

if that phase is not completed then the current payload design is ready for flight.

Payload Requirement Verification Final verification of payload requirements is accomplished through flight testing and post flight analysis. The functional requirements of the payload and the design feature that satisfies each requirement are compiled in Table 41.

SOW # Requirement Satisfying Design Feature Verification Method

3.1.3.1

UV Radiation SU-100

Field Testing, Static and Flight Testing, Analysis, Demonstration

Solar Irradiance TSL2561, SP-110

Humidity HIH4030

Temperature BMP180

Pressure BMP180

3.1.3.2

0.2Hz Data During Descent

16MHz Arduino Mega 2560, Software

Lab Testing, Flight Testing, Analysis

3.1.3.3

0.016Hz Data After Landing

16MHz Arduino Mega 2560, Software

Lab Testing, Flight Testing, Analysis

3.1.3.4

Post-Landing Data Termination

Software Lab Testing, Analysis

3.1.3.5

2 Descent Pictures

VC0706, Keyfob Video Camera

Static and Flight Testing 3 Landing Pictures

VC0706, Keyfob Video Camera

3.1.3.6

Horizon Orientation

ARTCOS Lab Testing, Analysis

3.1.3.7

Onboard Data Storage

Micro SD Lab Testing, Static and Flight Testing, Analysis

Data Transmission

900MHz XBee Radios Range Testing, Flight Testing

3.1.3.9

GPS LS20031 Lab Testing, Static and


118


SOW # Requirement Satisfying Design Feature Verification Method

Flight Testing, Analysis

3.2

Scientific Method Scheduling, Analysis, Testing, Documentation

Inspection

3.4

Jettisoned Components

No Separation, Clear Acrylic Housing

Inspection

3.5

Recoverable and Reusable

Aluminum Framework, No Separation

Lab Testing, Static and Flight Testing, Analysis

Verification Statements

The payload verification statements are listed below. The statement number is followed by the corresponding requirement from the SOW in parenthesis, followed by the statement with the results of the verification.

1. (3.1.3.1) UV radiation sensor functionality was verified in a field test, where accurate measurements were taken from within the acrylic housing structure. The solar irradiance, humidity, pressure, and temperature sensors‟ functionality has been verified in field testing and test flights.

2. (3.1.3.2) Telemetry data from test flights was analyzed to determine a data sampling rate of 0.83 Hz.

3. (3.1.3.3) The data sampling rate is programmed to collect measurements every 60 seconds after landing, and has been verified through lab and flight tests.

4. (3.1.3.4) The payload deactivates 10 minutes after detecting landing, and has been verified in the lab.

5. (3.1.3.5) The payload gathered the required number of photographs, as verified with the full scale test flight on 3/15/2013.

6. (3.1.3.6) Image orientation has been verified through lab testing. 7. (3.1.3.7) Data transmission and storage has been verified in lab and flight tests. 8. (3.1.3.9) The payload design incorporates a transmitting GPS unit. 9. (3.2) Scientific method is validated through copious amounts of detailed

documentation. 10. (3.4) The payload design can be verified to meet this requirement due to the fact

that it remains within the vehicle throughout the flight. 11. (3.5) The payload is recoverable and reusable, as verified through sub-scale and

full scale test flights.

Table 41: Payload Functional Requirements


119


Safety and Environment (Payload)

Failure Modes and Effects Analysis

Each failure mode for the payload has been assigned a Risk Priority Number (RPN). The RPN is a ranking system to determine which failure modes are more likely to occur, and require the most attention. The RPN value is calculated by taking the product of three scores: severity, occurrence, and detection. Severity is the importance of the effect upon the project. Severity is scored on a scale from one to five, where one is not severe and five is very severe. Occurrence is the frequency with which a given cause occurs and creates failure modes. Occurrence is scored on a scale from one to five, where one is not likely and five is very likely. Detection is the ability to detect or prevent a given cause. Detection is scored on a scale from one to five, where one is easy to detect and five is not easy to detect.

Payload Integration Failure Modes

Table 42 provides an updated summary of potential failure modes that could occur to the payload subsystem during integration, along with associated severity, occurrence, detection, and RPN.



Proposed Mitigation


Sev

erity

Occu

rren

ce

Dete

ctio

n

RP

N

Screw hole stripped out

Over tightened screws

Inadequately secured payload

Ensure the bulkhead is

replaced when required


3 1 1 3

Incompatible hardware

Lack of research

Payload will not integrate

properly

Ensure precision of fit

during manufacturing


3 1 1 3

Components damaged

during integration

Lack of attention

when integrating

payload components

Electronic malfunction

Be careful while inserting payload


5 1 1 5

Launch Operations Payload Failure Modes

Table 42: Potential Failure Modes for Payload Section During Integration


120


Table 43 provides an updated summary of potential failure modes that could occur to the payload subsystem during launch operations, along with associated severity, occurrence, detection, and RPN.



Proposed Mitigation


Sev

erity

Occu

rren

ce

Dete

ctio

n

RP

N

Battery disconnects

Improperly restrained

Partial or complete

failure of the SMD

payload

Secure battery

terminals


5 2 1 10

Port holes improperly sized

Lack of research

Inaccurate sensor

readings

Inaccurate sensor

readings


3 1 1 3

Radio dislodges from breakout

board


Telemetry lost and possible

damage to SMD

Payload

Secure the radio to the radio mount


5 2 1 10

BMP180 exposed to excessive sunlight

Excessive exposure to

direct sunlight

Inaccurate readings

Mount sensor in a covered

location


1 1 1 1

ARTCOS mounts break

Wrong epoxy used

Camera will not be

capable of orienting

properly and possible

damage to SMD

Payload

Ensure all components

have adequate

mounting and thoroughly inspected


5 2 1 10

Wired connections disconnect

Lack of secure

connection

Component and/or System Failure

Use of braided wire in non-static components and thorough

inspection


5 3 1 15

Component breaks away from board


Partial or complete

failure of the SMD

payload and possible

damage to SMD

Payload

Ensure all components

have adequate

mounting and are

thoroughly inspected


2 2 1 4


121




Proposed Mitigation


Sev

erity

Occu

rren

ce

Dete

ctio

n

RP

N

Battery power fails

Low battery voltage

Partial or complete

failure of the SMD

payload

Check battery voltages prior to every flight and replacing

batteries regularly


5 1 1 5

Circuits become un-mounted


Partial or complete

failure of the SMD

payload and possible

damage to SMD

Payload

Ensure all circuits have

adequate mounting and

are thoroughly inspected


2 2 1 4

Likelihood and Potential Consequences for the Top 10 Failures

Table 44 provides the top five failure modes for the vehicle based upon the highest RPN value in ascending order, and their associated consequences.

RP

N


15 Wired connections disconnect Lack of secure

connection Component and/or

System Failure

10 Battery disconnects Improperly restrained

Partial or complete failure of the SMD

payload

10 Radio dislodges from breakout

board Improperly restrained

Telemetry lost and possible damage to

SMD Payload

10 ARTCOS mounts break Wrong epoxy used

Camera will not be capable of orienting

properly and possible damage to SMD

Payload

5 Battery power fails Low battery voltage Partial or complete failure of the SMD

payload

Table 43: Potential Failure Modes for Payload During Launch Operations Table XX Potential Payload Failure Modes during Launch

Table 44: Top Five Potential Failure Modes by RPN


122


Personnel Hazards

Potential hazards to personnel through the course of the project are provided in Table 45. 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.


Laceration

Knives, routers, saws,

file, dremel tool

Medium Serious injury

or death



operators manual

Discontinue all

operations, apply first

aid, contact EMS

Burns

Chemicals (FFFFg,

fiberglass resin),

welders, soldering Iron

Medium Minor to serious

injury



operators manual

Discontinue all


aid, contact EMS

Respiratory Damage

Chemicals (epoxy, solder), fumes,

fiberglass

Low Brain damage

or death


tool and equipment use, personal safety attire, consult

MSDS

Discontinue all


aid, contact EMS

Vision Damage

Welders, fiberglass, grinders, projectile

debris

Low Partial to complete blindness

Use of goggles, force shields,

consult MSDS, first aid kit

available, refer to operators

manual

Discontinue all


aid, contact EMS, use eyewash

Allergic Reaction

Epoxy, chemicals, fiberglass

Low

Loss of respiration,

inflammation (Internal & External)

Use of gloves, consult MSDS,

first aid kit available

Discontinue all


aid, contact EMS,

administer antihistamine

s, safety shower


123



Hearing Damage

FFFFg, Grinders, Ignition, Routers

Low Partial to complete deafness

Ear muffs, consult MSDS,

first aid kit available, refer

to operators manual

Discontinue all


aid, contact EMS

Dismemberment

Projectiles, Saws,

Launches Low

Permanent injury or death

Make sure proper safety measures are

taken, operators manual

Discontinue all

operations, apply first aid, and

contact EMS, tourniquet

The material safety data sheet (MSDS) that the manufacturer provides contains information about the material in consideration. It is comprised of 16 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. According to the safety plan, a binder containing all the MSDSs is always made available for personnel and brought to every launch. 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 component used during the project are kept in an operator manual 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 five 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. The team has obtained flight waivers for full scale launches.


There are no environmental concerns for the payload that are not also inherent to the vehicle. This is because the payload does not detach from housing during descent.

Table 45: Potential Hazards to Personnel


124


V) Launch Operations Procedures

Checklists

Recovery Preparation Checklist (to be completed before arriving to launch site):

1. Open altimeter software on dedicated GS computer, power on altimeters and

connect via USB

2. Calibrate all altimeters and verify they are functional for flight.

3. Run each altimeter in flight simulation mode to further verify functionality. This

includes using a multimeter to measure applied current and voltage from each

altimeter at the time of each deployment event.

4. Check battery voltage for each avionics bay, replace if below 5V (brand new

batteries will be used for the official flight).

5. Disable all unused outputs.

6. For Stratologger primary on the main parachute deployment, set to 700 ft AGL.

7. For Stratologger backup on the main parachute deployment, set to 650 ft AGL

(i.e. 50 ft delay)

8. For Stratologger primary on the drogue deployment, set to deploy at apogee

9. For Stratologger backup on the drogue deployment, set 1 second delay from

apogee.

10. Clear all unused banks.

11. Screenshot altimeter configuration and parameter input page for documentation.

12. Inspect physical wiring configuration of altimeters.

13. Inspect altimeter mounting, battery mounting, and electrical connection mounting.

14. Ensure proper orientation and load avionics sleds into couplers.

15. Weigh Dog barf for each section

16. Verify that Astro DC40 is charged.

17. Astro DC40 is secured in nosecone. Verify that it is not powered before replacing

nose cone.

18. Inspect main parachute and shroud lines for physical damage.

19. Inspect main parachute deployment bag and pilot drogue for physical damage.

20. Inspect main parachute shock chord (40 ft) for physical damage (brand new

shock chord will be used the day of the official launch)

21. Inspect drogue parachute shock chord (20 ft) for physical damage (brand new

shock chords will be used the day of the official launch)

22. Inspect drogue parachute and shroud lines for physical damage.

23. In


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24. Pack main parachute in deployment bag.

25. Store main parachute in deployment bag with 40 foot shock chord length.

26. Pack drogue parachute in nomex cloth.

27. Store drogue parachute with 20 foot shock chord length.

28. Prep all ejection canister leads but cutting them to length ( 3 foot for main

parachute bay, 2 foot for drogue parachute bay) and ensure no exposed wire

along the length of the lead.

29. Test continuity of ejection canisters leads.

30. Attach shorted quick connect to ejection canisters leads. This prevents static

buildup on the ematch and reduces the risk for inadvertent charge detonation.

31. Verify extension wires are routed through both avionics bay coupler lids (drogue

and main).

32. Seal lead holes with all-purpose putty.

33. Secure avionics lids.

34. Weigh vehicle and record total weights.

35. Update simulation with total weights, get predicted flight performance, and

calculate any needed ballast. Add ballast to nose cone and re-weigh if needed.

36. Document simulation model for post flight comparison and analysis.

37. Pack tool boxes and verify that inventory is ready to be taken to launch site. This

includes all backup hardware that maybe needed on the field:

a. Gray tool box

b. Black tool box

c. Black and Green tool box

d. Nose cone (ballast system, GPS) attached to upper body airframe

e. Payload housing structure and couplers/avionics sleds and bays attached

and enclosed.

f. Booster section with motor casing and retaining ring

g. Ground Station Computer

h. Canopy

i. Folding table

j. Chairs

Motor Preparation Checklist:

1. Ensure that the following components are ready for motor assembly:

a. Motor Grains (box of 3)

b. Grain spacer O-rings

c. Smoker Tracker Charge


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d. Smoke Tracking Insulator

e. Nozzle Holder

f. Nozzle

g. Nozzle O-ring

h. Phenolic Case Liner

i. Forward Closure

j. Forward Insulating Disk

k. Retaining Ring

l. Motor Wrench

m. Motor Housing

n. Motor Reload Kit

o. Igniter

p. Grease

q. Gloves

2. Follow the motor assembly instructions carefully.

3. Confirm motor is ready to be loaded into rocket.

4. Give Jake, the lead engineer, the igniter for installation later.

Igniter Installation Checklist:

1. Inspect igniter for any physical damage.

2. Test continuity of igniter with multimeter.

3. Insert igniter into motor until firmly seated against the smoke tracker.

4. Twist a loop in igniter wire to ensure no igniter movement.

5. Replace motor nozzle cap.

6. Touch leads from ignition system to ensure no static buildup.

7. Securely connect ignition box leads to igniter wires by wrapping each alligator

clip completely.

Launchpad Setup Checklist:

1. Confirm that vehicle is ready to be loaded on launch pad.

2. Upon clearance from RSO, take rocket to designated launch rail.

3. Test ignition box continuity (n/a for Hunstville launch site).

4. Lower launch rail.

5. Inspect launch rail for obstructions and rail buttons.

6. Set launch rail to specified angle and measure to confirm.

7. Externally arm avionics and payload with magnet.

8. Confirm altimeters are powered via LED indicator and armed via output beeps.


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9. Confirm payload is activated and sending telemetry. Verify ARTCOS starting

orientation.

10. Everyone except the lead engineer leaves the launch pad area.

11. Begin Igniter install checklist.

Launch Procedures Checklist:

1. Call in flight and confirm flight waiver at 817-858-7527(n/a for Huntsville).

2. Initiate motor preparation.

3. Initiate payload preparation and ground station setup.

4. Check all port holes for obstructions (acrylic payload housing and static

fiberglass port hole rings)

5. Weigh black powder and pack each ejection canister with wooden dowel rod and

dog barf wadding. The weights are as follows:

a. Main Parachute Deployment charge (primary)= 4.8 grams

b. Main Parachute Deployment charge (secondary)= 5.2 grams

c. Drogue Parachute Deployment charge (primary)= 2.6 grams

d. Drogue Parachute Deployment charge (secondary)= 2.8 grams

6. Upon Completion of Motor Prep, load motor and screw in retaining ring.

7. Attach drogue shock chord to motor eyebolt.

8. Remove nose cone and activate GPS tracker. Verify that a link has been

established between receiver and transmitter.

9. Replace and attach nose cone.

Upon completion of Payload and Ground Station prep, payload housing with payload is

ready to be loaded with black powder charges, recovery parachutes and attachment

harnesses.

10. Await electronics team to bring flight ready payload in acrylic housing structure to

recovery team.

11. Disconnect shorted quick connect from ejection canister leads.

12. Connect ejection charge leads from avionics lid (labeled A) to canisters for

booster section.

13. Run the ejection charges into the bottom of the booster section.

14. Pack recovery wadding into booster section.

15. Z-fold drogue shock chord and connect to coupler A eyebolt.

16. Attach drogue parachute and wrap in nomex.

17. Pack drogue shock chord into booster, followed by drogue wrapped in nomex.

18. Connect booster section to payload section via coupler A.

19. Attach main parachute and deployment bag (with pilot drogue):


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a. Attach shock chord to coupler B eyebolt.

b. Attach rear deployment bag line to pilot drogue.

c. Attach main parachute shroud lines to upper body airframe attachment

eyebolt.

d. Verify forward deployment bag line (towards flap) is attached to main

parachute.

e. Attach shock chord to upper body airframe attachment eyebolt.

20. Disconnect shorted quick connect from ejection canister leads

21. Connect ejection charge leads from avionics lid (labeled B) to ejection canisters

for upper body airframe section.

22. Run the ejection charges into nose cone section.

23. Pack recovery wadding into nose cone section.

24. Pack deployment bag. The linear packing order from nose cone to booster is as

follows:

a. Z-folded shock chord

b. Deployment bag with flap facing towards nosecone.

c. Pilot drogue.

25. Connect nose cone section to payload housing structure via coupler B.

26. Insert sheer pins between upper body airframe and payload housing structure.

27. Begin Setup on Launchpad Checklist.

28. Confirm that Launchpad setup and igniter installation checklists are complete.

29. Arm ignition system box (n/a for Hunstville).

30. Exit launch rail area.

31. Initiate countdown sequence/ await LCO to announce and initiate launch.

(Hunstville).

32. 5…4…3…2…1…Fire!

33. Observe flight carefully, listening for black powder reports and visually confirming

the descent and staging of recovery devices.

34. Track vehicle with Garmin Astro GPS.

35. Upon clearance from RSO, move to recover vehicle from field.

36. Begin Post Flight Inspection Checklist.

Troubleshooting Checklist:

Troubleshooting procedures are implemented upon two main areas: motor ignition failure and electronics. For motor ignition failure,

1. Remove igniter from motor and replace nozzle cap.

2. Test continuity of igniter.


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3. Verify ignition system continuity and ensure adequate power is being supplied to

igniter.

For electronics, 1. If payload electronics or avionic altimeters are not functioning properly, or

telemetry is not being received at the ground station, using magnet to disable

recovery electronics and power down payload.

2. Disconnect igniter from ignition system leads and remove igniter from motor.

3. Lower launch rail and remove vehicle.

4. Return to ground station, and disassemble vehicle.

5. Open avionics bays and inspect.

6. Remove payload and inspect.

Post-flight Inspection Checklist:

1. Do not disturb the landing site until photographic documentation has been

gathered.

2. Listen to the official altimeter to report achieved apogee. This requires disabling

the recovery avionics with magnet.

3. Disable payload electronics.

4. When given clearance, begin inspection.

5. Inspect vehicle for damage.

6. Inspect parachutes and attachment hardware for physical damage.

7. Inspect payload for damage.

8. When the official gives clearance, pack up and clear the landing site.

9. Report official altitude to NASA judge.

10. Return to HQ

11. Unload motor and thoroughly clean for next launch.

12. Remove GPS tracking device and connect to charger.

13. Complete post flight analysis and write up. This includes comparing predictions

to actual flight performance, safety and failure analysis, watching video of launch,

reviewing photographs from landing site, and updating the Facebook page with

new videos or images.

Payload Preparation (to be completed before arriving at launch field):

1. Charge Keyfob batteries 2. Test complete circuit and verify functionality. Test and install brand new batteries. 3. Verify telemetry accuracy. 4. Test GUI and Yagi antenna.


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5. Check microSD cards for data (ADGS, ARTCOS, Keyfob video). 6. Verify data accuracy. 7. Verify attachment hardware. This includes battery retention, circuit board mounts

and standoffs. 8. Verify cable management and snap connectors. 9. Verify payload framework attachment to bulkhead. 10. Pack payload and acrylic housing structure 11. Verify brand new batteries and tie wraps are packed. 12. Verify Yagi, Xbee receiver, USB interface cable and Ground Station Computer

are packed.

Pre-launch Payload Preparation and Ground Station Setup:

1. Setup Ground Station table. 2. Setup Ground Station computer. 3. Activate Payload. 4. Remove SU and SP lens covers. 5. Verify Telemetry and Ground Station GUI. 6. Check battery voltage and change if necessary. 7. Deactivate Payload 8. Turn on Keyfob. 9. Secure official altimeter. 10. Load payload into acrylic. 11. Secure bulkhead screws. 12. Bring payload housing structure with SMD ready to recovery team.

Safety Materials Checklist

1. MSDS Binder 2. Operators Manuals Binder 3. Launch Procedures Binder 4. First Aid Kit 5. New launch operations checklists on clipboard. 6. Fire extinguisher

Safety Checklist

This checklist ensures that the instructions from the awarded flight waivers have been followed and the team is in accordance with the law.

1. The Lockheed Martin Fort Worth Flight Service Station has been contacted and a Notice to Airmen has been issued.


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2. The Fort Worth ARTC Stephenville Supervisor has been contacted the day of the scheduled launch.

3. All applicable operating limitations of Federal Aviation Regulations (FAR‟s), Title 14, Part 101, expect for parts (e,f,g) have been checked.

4. Ensure that operations will be conducted in accordance with all applicable state and local ordinances.

5. http://www.aviationweather.gov/adds/metars/ has been checked for horizontal visibility of more than five miles and cloud coverage is less than five tenths of the intended altitude of the vehicle.

6. Aircraft spotters have been assigned and informed of their job. 7. Make sure that a safe launch radius has been obtained before launch (300 feet

or greater) 8. A list of authorized personnel for launch operations has been made. 9. The team has been briefed on the following, expected altitude, horizontal

visibility, and cloud coverage height, who the aircraft spotters are, what members are allowed in the flight operation area, and what the simulated landing radius of the vehicle will be given the different wind speeds.

The following is an itemized components checklist for each subsystem.

Structure Components:

1. Nose Cone 2. Upper body airframe 3. Booster Section with retaining ring 4. Avionics Bay A (shielding, avionics sled, bay lid and attachment hardware,

welded eyebolt) 5. Avionics Bay B (shielding, avionics sled, bay lid and attachment hardware,

welded eyebolt) 6. Payload Housing Structure 7. Payload Framework 8. Payload Framework bulkheads and attachment hardware 9. Motor Housing with welded eyebolt 10. Motor Reload Kit 11. Motor Grains

Recovery Components:

1. Black Powder (3F) 2. Ejection Canisters (x4) 3. Dowel Rod (for packing charges) 4. 9V battery (x2) 5. Gorilla Tape

http://www.aviationweather.gov/adds/metars/


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6. Stratologger Altimeters (x4) 7. Altimeter Interface Cable 8. Magnet 9. Main parachute 10. Main parachute deployment bag 11. Pilot drogue 12. Drogue parachute 13. Nomex protector cloth 14. Recovery wadding (Dog Barf) 15. Main Shock Cord 16. Drogue Shock Cord 17. Garmin GPS tracker (handheld receiver and transmitter) 18. Garmin GPS Box with DC and AC chargers 19. Nylon shear pins

Payload and Ground Station Components:

1. Yagi antenna 2. Xbee interface cable 3. Xbee receiver 4. Stratologger programming software 5. Payload programming software 6. GUI 7. Ground station computer.

Safety and Quality Assurance

Risk Assurance

Based upon the RPN values calculated for each considered risk or failure mode previously in the Safety and Environment sections, it can be concluded that all risks are within acceptable levels. The greatest theoretical RPN would be a rating of 5 for severity, occurrence, and detection, resulting in a RPN= 125. The highest recorded RPN for any single risk considered is 25 or 20%.

Risk Assessment for Launch Operations

Table 46 is a risk assessment for launch operations with proposed mitigations and completed mitigations.


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Category Risk Consequences Completed Mitigations

Recovery Preparation

Premature black power

canister ignition

Explosion, personnel injury, possible damage to vehicle, and components

damaged

Testing, recovery altimeter shielding, ensure e-match leads are grounded, , e-matches armed before

launch only, and procedural checklists followed

Main parachute

catches wind

Personnel Injury, parachute could become

snagged and tear.

Ensure that the main parachute is properly

inserted into deployment bag, place in storage until

needed, and follow procedural checklists

Main or drogue

parachute folded

incorrectly

Parachute either becomes entangled upon

deployment or is not able to fully envelope in flight

Follow manufacturer instructions, and follow procedural checklists

Altimeters incorrectly

wired

Ballistic descent, vehicle components damaged,

and personnel injury

Follow wiring schematic, recovery lead checks wiring,

and follow procedural checklists

Shock cords not attached to eye hooks

or parachutes

Ballistic descent, and personnel injury

Follow procedural checklists

Motor Preparation

Improperly assembled

motor

Explosion on launch pad causing injury to personnel and vehicle components

Follow manufacturer instructions, and procedural

checklists

Assembled motor

dropped or improperly

loaded

Explosion on launch pad causing injury to personnel and vehicle components

Do not use motor, assemble another motor, properly store, and load motor in safe controlled

environment

Igniter installation

Motor does not ignite

Rocket does not launch Check igniter for continuity,

bring spare igniters

Setup on Launch Pad

Launch rail slot jammed

Unstable launch, unpredictable flight path,

personnel injury

Inspect launch rail before vehicle is put on rail , and

follow procedural checklists

Unarmed avionics bays

Ballistic descent, personnel injury, and

damage to vehicle components

Swipe magnets across magnetic switches, and



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Category Risk Consequences Completed Mitigations

Payload unarmed

No flight data, scoring altimeter inoperable, loss of points at competition

Swipe magnets across magnetic switches, and


Launch Procedures

Failure to Launch

Rocket not launched, payload, and avionics bay

record no flight data Follow procedural checklists

Trouble Shooting Delay in launch

operations

Allotted preparation time expires

Follow procedural checklists, rehearsed

preparation

Post-flight inspection

Walking to recover vehicle

Possible personnel injury from falling or wildlife

Launch area be checked out before launch , and recovery team will be alert and focus

on their surroundings

A black powder

canister not being set off

in flight

Possible explosion and personnel injuries

If discovered recovery team member will not move, until tools have arrived so that it can be safely deposed of


Please refer to environmental concerns in the Safety and Environment (Vehicle) Section.

Identification of person responsible

The team safety officer, Blake, is level one certified with NAR, and has obtained a FAA flight waiver in his name for the use of full scale launches. 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. .

Table 46: Risk Assessment for Launch Operations


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VI) Project Plan

VI) Project Plan

Budget Summary The following chart includes the 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 a budget that is sufficient in meeting all costs necessary to complete the mission. Table 47 breaks down the known project costs. The team has allocated funds to several areas of the project as charted in Table 48, 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

Element Cost Testing/Prototyping $13,972.23

Outreach $3,669.77

Final Build $4,007.21

Travel to Competition $8,200.00

Total $29,849.21

In the following charts, itemized budgets are compiled to illustrate an “on the pad” cost for the final build design. The data from Tables 49-52 are broken down between three categories. The first is Structure and Propulsion, the second is Recovery, and the last is SMD Payload, which has been subdivided between through-hole, PCB, and Non PCB budgets.

Table 47: Known Project Costs

Table 48: Team Budget


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VI) Project Plan

Body Part Number Price Per

Unit Quantity Total

Nose Cone FNC5.5EL $47.50 1 $47.50

Fiberglass Body Tube

G12-5.5-60 $38.84 5.3 $233.04

Acrylic Body Tube

ACRCAT5.500ODX.250 $39.05 1 $39.05

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

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

Bulk Plate (Payload)

PVCGRAY2.00LAM12x24 $147.82 0.25 $36.96

Bulk Plate Standard

FBP5.5 $7.60 3 $38.00

Couplers G12CT-5.5 $47.03 1 $47.03

Centering Rings

FCR5.5-3.0 $8.55 4 $34.20

Epoxy 4500Q $69.00 0.25 $17.25

Motor Retainer

RA75 $52.00 1 $52.00

Casing and Hardware

Cesaroni 3 Grain Case and closures

260.96 1 260.96

Motor Reload Cesaroni L-1720 170.96 1 170.96

Miscellaneous Hardware $100.00 1 $100.00

Total $1189.61

Projected

Total $1199.64

Table 49: Structure/Propulsion System Budget


137

VI) Project Plan

Proposed Selection

Item Number Unit

Cost

Quantity Typical

Unit

Cost

Altimeters StratoLogger $79.95 4 $319.80

Electric Matches XL Variable-Capacity Ejection Canister

$2.75 4 $11.00

FFFg Black Powder Goex 3F Black Powder $3 1 $3

Main Shock Cord Tubular Kevlar $37.99 1 $37.99

Drogue Shock Cord Tubular Kevlar $31.49 1 $31.49

Main Parachute 10ft Cert 3 $239.00 1 $239.00

Flameproof Main Parachute Deployment Bag

DB8 $40.00 1 $40.00

Drogue Parachute 2ft $27.50 1 $27.50

Flame-Proof Drogue Parachute Nomex

18x18 $10.95 1 $10.95

Eye bolts 0.25in. (compact) $2.00 2 $4.00

Quick Links 0.25in. Stainless Steel Delta Quick Link

$2.99 6 $17.94

Shear Pins 2-56 Nylon Shear-Pin (10 pack)

$1.00 1 $1.00

Arming Switches Featherweight Magnetic Switch

$25.00 2 $50.00

GPS Garmin DC-40 $144.50 1 $144.50

Battery Ultralife U9VLBP $6.65 2 $13.30

Total $951.47

Projected $1383.35

Table 50: Recovery System Budget


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VI) Project Plan

Price Price Quantity Total

Arduino 2560-R3 $58.95 1 $58.95

Arduino Pro Mini 328 - 5V $18.95 2 $37.90

Micro SDHC Card $9.99 3 $29.97

Adafruit 254 – Micro SD Adapter $15.00 2 $30.00

BMP180 – Pressure/Temperature Sensor $15.00 1 $15.00

HIH4030 – Humidity Sensor $16.95 1 $16.95

SP-110 – Pyranometer $169.00 4 $676.00

TSL2561 – Lux Sensor $12.50 1 $12.50

SU-100 – UV Sensor $159.00 4 $636.00

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

Adafruit 397 – Camera $42.00 1 $42.00

Key fob 808 #20 $20.00 1 $20.00

Xbee-PRO XSC S3B $42.00 1 $42.00

XBee 0.1” Through Hole Spacing Adapter $2.99 1 $2.99

3.5” RPSMA 900MHz Antenna $14.64 1 $14.64

LS20031 GPS $60.00 1 $60.00

9V Battery Holder $2.95 4 $11.8

LCD – Sparkfun 11062 $34.95 1 $34.95

ADXL345 – Accelerometer $27.95 1 $27.95

Ultralife U9VLBP - 9V Battery $6.65 4 $26.60

Adept A1E Altimeter $29.95 1 $29.95

Total $1,866.13

Projected

Total $1558.24

Table 51: Payload Budget (Through-Hole PCB)


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VI) Project Plan

Product Price Quantity Total

ATMega2560 $17.97 1 $17.97

ATMega16U2 $3.71 1 $3.71

Arduino Pro Mini 328 - 5V $18.95 2 $37.90

MicroSD Card Connector $3.08 1 $3.08

Adafruit 254 – Micro SD Adapter $15.00 1 $15.00

BMP180 (Chip Only) – Pressure/Temperature Sensor $2.88 1 $2.88

HIH4030 (Chip Only) – Humidity Sensor $13.68 1 $13.68

SP-110 – Pyranometer $169.00 2 $338.00

TSL2561 (Chip Only) – Lux Sensor $2.84 1 $2.84

SU-100 – UV Sensor $159.00 2 $318.00

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

Adafruit 397 – Camera $42.00 1 $42.00

HackHD - 1080p Camera Module $159.95 1 $159.95

Xbee-PRO XSC S3B $42.00 1 $42.00

XBee 0.1” Through Hole Spacing Adapter $2.99 1 $2.99

3.5” RPSMA 900MHz Antenna $14.64 1 $14.64

LS20031 GPS $60.00 1 $60.00

9V Battery Holder $2.95 3 $5.85

LCD – Sparkfun 11062 $34.95 1 $34.95

ADXL345 – Accelerometer $27.95 1 $27.95

Tenergy Li-Ion 3.7V Battery $9.90 1 $9.50

Ultralife U9VLBP - 9V Battery $6.65 3 $53.20

Adafruit Buck Converter $14.95 2 $28.90


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VI) Project Plan

Adept A1E Altimeter $29.95 1 $29.95

47 µF Capacitor $2.25 2 $4.50

100 nF Capacitor $0.19 12 $2.28

22 pF Capacitor $0.35 3 $1.05

1 µF Capacitor $0.29 3 $0.87

100 nF Polarized Capacitor $0.77 2 $1.54

10 nF Capacitor $0.25 2 $0.50

10 µF Capacitor $0.42 1 $0.42

2.2 µF Capacitor $0.96 1 $0.96

10 µF Polarized Capacitor $0.43 2 $0.86

Switching Diode $0.06 1 $0.06

Resettable Fuse $0.46 1 $0.46

EMI Filter Bead $0.40 1 $0.40

215 nH Inductor $1.90 1 $1.90

Green LED $0.14 1 $0.14

Red LED $0.08 1 $0.08

White LED $0.17 2 $0.34

Yellow LED $0.08 1 $0.08

Blue LED $0.11 1 $0.11

10 kΩ Resistor $0.25 3 $0.75

1 kΩ Resistor $0.81 9 $7.29

1 MΩ Resistor $0.20 2 $0.40

22 Ω Resistor $0.19 2 $0.38

330 Ω Resistor $0.20 3 $0.60

68 kΩ Resistor $0.20 1 $0.20


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VI) Project Plan

4.7 kΩ Resistor $0.19 2 $0.38

500 kΩ Resistor $0.55 1 $0.55

50 kΩ Resistor $1.40 1 $1.40

25 Ω Resistor $0.58 4 $2.32

5 kΩ Resistor $0.08 4 $0.32

250 Ω Resistor $0.45 4 $1.80

Plastic Spacers $0.19 4 $0.76

P-Channel MOSFET $0.43 1 $0.43

16 MHz Crystal $0.96 1 $0.96

16 MHz Crystal $1.16 1 $1.16

ESD Suppressor Diode $0.25 2 $0.50

Operational Amplifier $0.83 1 $0.83

Operational Amplifier $0.67 1 $0.67

Line Driver $0.38 1 $0.38

Digital Potentiometer $2.06 1 $2.06

USB Connector $1.07 1 $1.07

Non-PCB $1346.68

PCB $85.92

Total 1432.6

Projected

Total 1558.24

The purchase of new items is shown in the bottom row. The bulk of the cost in building the PCB design comes from the costly elements that cannot be reproduced (SP-110, SU-100, etc.).

Table 52: Payload Budget (PCB/Non PCB)


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VI) Project Plan

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 (as seen in Figure 88), 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.

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 the Gantt Testing Timeline in Figure 89. The following chart gives a visual representation of major project deliverables.

Figure 88: President's Circle Awarding the Aeronautical Team


143

VI) Project Plan


144

VI) Project Plan


145

VI) Project Plan

Figure 89: Gantt Testing Timeline


146

VI) Project Plan

Outreach Timeline

The chart below delineates the dates of the team‟s educational outreach events. 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 gave 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 and Figure 90.

Education plan

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

Figure 90: Outreach Timeline


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VI) Project Plan

subjects with the intent of increasing the number of people that choose to pursue STEM related careers.

Educational Outreach

Educational outreach targets students, teachers, and the community as a whole through a variety of events. 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 receive an opportunity to work closely with the team members on a lesson which reinforces concepts learned previously. 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.

Figure 91: Students Launching a Water Rocket


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VI) Project Plan

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 that 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.

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.


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Accomplished Educational Outreach

October 5, 2012

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 53 and illustrated in Figure 93.

Figure 92: Team Members Educate and Entertain Acton Students


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23%

20%

21%

19%

8% 9%

Subject Interest

Science Math Engineering Technology Rocketry None

Subject Interest Count Science 39

Math 35

Engineering 36

Technology 33

Rocketry 14

None 15

Table 53: Educational Outreach Survey Results

Figure 93: Educational Outreach Survey Results


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3% 13%

5%

19%

13% 8%

15%

9% 3%

7% 5%

Presentation Learning Outcomes

Forces Propulsion Construction Design

Procedures Failure Modes Qualifications Careers

Competitions NASA None

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 54 and illustrated in Figure 94.

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 54: Presentation Learning Outcomes

Figure 94: Presentation Learning Outcomes


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13%

4%

6%

12%

32%

25%

8%

Favorite Part

Launch Launch Failure Flight

Outside Water Rockets Big Rocket

None

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 55 and illustrated in Figure 95.

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.

November 9, 2012

Favorite Part Count Launch 11

Launch Failure 3

Flight 5

Outside 10

Water Rockets 27

Big Rocket 21

None 7 Table 55: Students' Favorite Portion of Presentations

Figure 95: Students' Favorite Portion of Presentations


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VI) Project Plan

The Tarleton Aeronautical Team co-hosted the Tarleton Star Party this semester. Before the sun set, the team presented information concerning the STEM fields and the necessity for growth and interest in those fields to the community and grade school students. The team discussed the importance of doing well in all levels of school as the knowledge compounds as students move from grade level to grade level. The NASA USLI competition and our involvement were also heavily discussed. Event participants were given time for a question and answering session as we watched team members launch off a series of rockets before the sun set. Although the setting didn‟t allow for a thorough educational evaluation, the team did receive many oral critiques and feedback, and the event was said to deliver a very positive and informative message. A few parents in the group mentioned that they better understood the importance of the STEM classes and they would be working more diligently with their children to encourage growth and understanding in their students‟ science and math classes. Other participants wanted more information on the competition and what was expected of the group. Two of the children present expressed the desire to have our group visit their school and one of the adults present said she came to our event after watching our presentation to the junior high earlier that day.

November 12, 2012

The Tarleton Aeronautical Team invited Glen Rose High School to come to Tarleton State University in Stephenville, Texas to attend an educational outreach. Students were divided up into three random groups upon arrival. Identical materials were distributed to each team and they were given a 20 minute time limit to create a water bottle rocket that would be flight ready. The group did a debriefing session following the exercise. During the debriefing, the team discussed important aspects that were called upon during the exercise. These included communication skills, team work, and the ability to recall prior STEM topics. A discussion about STEM fields occurred. This was followed by a question and answering session, where the students were able to question our team about STEM fields, the competition, college, and future careers. An evaluation of the educational engagement proved that the outreach was highly valuable and influential. Students gave positive feedback about the event. It was stated that it was very informative and interesting. Their teacher appreciated the impact of the impromptu rocket building session as an introduction to the importance of the STEM fields. During the session, the team paid close attention to the classes the students were taking, in order to make the information relevant to what the students were already learning. This correlation helped the students understand the importance and relevance


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VI) Project Plan

of STEM to their lives and their education. Furthermore, one of the students decided after visiting with us and our school to pursue a degree in computer science when he graduates high school.

December 17, 2012

On December 17, 2012, members of the Tarleton Aeronautical Team held an outreach at Morgan Mill ISD in Morgan Mill, TX. The team hosted a “rocket fair” for the students. Students in grades 4-8 attended this portion of the event. The students rotated through a series of six stations where Tarleton Aeronautical Team members used a variety of formats to discuss STEM topics. The stations are discribed in Table 56 below.

Station Description

History The history of space exploration & NASA were investigated at this table.

Vocabulary Word searches, matching, and an introduction to rocketry related words. A small rocket was used as a visual for this station.

Math

Discovering and calculating speed & velocity. Students were able to walk and find the distance, time and direction of their travel. They then learned how to calculate speed & velocity. Discussions about the correlation to rocketry occurred.

Technology Video presentations of rocket launches. Discussion about STEM fields and their importance. Discussion about the USLI competition

Problem Solving

Wind tunnel discussion. Students cut paper cups to experiment with the concepts of stability and lift. To test their designs, students placed their creations in a team created wind tunnel. Results were discussed as well as STEM implications.


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VI) Project Plan

The team conducted a pre-test before the students participated in the stations and a post-test consisting of six questions. Each question related to the content discussed during the six various stations in the rocket fair. The difference in the student‟s responses from pre-test to post-test was quite exciting. The group, as a whole, showed growth in knowledge gained. Questions that were commonly left unanswered on the pre-test were correctly answered following the rocket fair.

Art

Students were provided coffee filter parachutes so that they could learn to fold a parachute as our team does before each launch. A parachute from our sub-scale rocket is used for the demonstration. Students were then able to create their own design for their parachute. Students learned about the recovery systems of our competition rocket. Duel deployment was discussed as well as how the STEM fields were used in conjunction with choosing our parachutes for the competition.

Table 56: Educational Outreach Stations

Figure 96: Students Received Stickers for Correct Responses


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VI) Project Plan

On the pre-test, students were prompted with these questions, “What does NASA stand for?”, and “What does STEM stand for?” The students‟ responses were all a variation on the team‟s favorites, “National Asteroid and Space Association” and “Space Team Economic Mission”. When pretesting, most students incorrectly answered these questions or chose to leave them blank. However, on the post-tests nearly every student could accurately identify what STEM stands for, and over half could identify the entire NASA acronym. Though some students did not correctly answer what NASA stands for, many did list a variety of facts they learned about NASA and thus showed a growth in knowledge. Beyond paper testing, a group discussion was conducted concerning the topics presented at the stations. Once started, the discussions quickly became student-led and had to be stopped and redirected by our team in order to cover all the topics in the allotted time. The Tarleton Aeronautical Team truly felt this was a success because it

Figure 97: Students Engaged in Learning


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VI) Project Plan

demonstrated the amount of information the students had received and the interest they found in the topics!

The verbal feedback from the teachers and school principle was highly positive and there was interest expressed in the team returning to the school again. After the rocket fair, the entire school (grades K-8) joined the team outside for a rocket launch. Discussions about the STEM fields continued, and students were given the opportunity to ask questions related to our visit. Students were also given a demonstration about how to properly prepare a rocket for launch. Following our event, the team was invited to stay for lunch and recess. Members of our group were able to talk in a less formal environment with students in grades 4-8 as they enjoyed their lunch and recess break. Photos of the event are available on the team‟s Facebook page under the “Morgan Mill ISD Outreach” album.

Figure 98: Team Members Addressed Large Audiences of Students

https://www.facebook.com/TarletonAeronauticalTeam/photos_albums


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VI) Project Plan

December 18, 2012

The Tarleton Aeronautical Team held an educational outreach event at Bluff Dale ISD in Bluff Dale, TX on December 18, 2012. The team hosted a “rocket fair” for the students. Initially it was planned for grades four through eight, but upon arriving at the school grades 2-4 were added to the group. Students rotated through a series of seven stations where Tarleton Aeronautical Team members discussed STEM topics in a variety of ways. The stations are described in Table 57.

Station Description

History The history of space exploration & NASA were investigated at this table.

Vocabulary Word searches, matching, and an introduction to rocketry related words. A small rocket was used as a visual for this station.

Math

Discovering and calculating speed & velocity. Students were able to walk and find the distance, time and direction of their travel. They then learned how to calculate speed & velocity. Discussions about the correlation to rocketry occurred.

Technology Video presentations of rocket launches. Discussion about STEM fields and their importance. Discussion about the USLI competition

Problem Solving

Wind tunnel discussion. Students cut paper cups to experiment with the concepts of stability and lift. To test their designs, students placed their creations in a team created wind tunnel. Results were discussed as well as STEM implications.

Art Students were able to create their own design for their parachute. Further discussion concerning STEM fields and the NASA competition were conducted.

Recovery

(This table was split from the art station into a new station due to the large number of students attending the event.) Students learned about the recovery systems of our competition rocket. Duel deployment was discussed as well as how the STEM fields were used in conjunction with choosing our parachutes for the competition. Students were provided coffee filter parachutes so that they could learn to fold a parachute as our team does before each launch. A parachute from our sub-scale rocket is used for demonstration.

Table 57: Educational Outreach Stations


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VI) Project Plan

Students in grades five through eight were given a pre-test and post-test at the outreach session. Students showed a large amount of improvement in their knowledge concerning NASA and the STEM fields. The pre-test and post-test asked the same questions and some students who could not answer any questions prior to the outreach, answered all questions correctly on the post-test. The final question on the test asks the student to list words that come to mind when they think of rockets. The most popular responses on the pretests are “fire,” “space,” and “loud” as a response. Tests returned after the rocket fair included words such as “launch,” “airframe,” “apogee,” and “Tarleton.”

The students in grades 2-4 were verbally evaluated, and they too demonstrated a growth in understanding of the topics presented during the team‟s visit. For example, many students did not know what NASA or STEM stood for, but could explain them after attending the rocket fair. Verbal feedback from students indicated excitement about rocketry and NASA, as well as an interest in science. The most mentioned station was the wind tunnel, the parachute folding, and the parachute design tables.

After the rocket fair, the entire school (grades K-8) joined the team outside for a rocket launch. Discussions about the STEM fields continued and students were given the opportunity to ask questions related to our visit. Photos of the event are available on the team‟s Facebook page. Visit in the album entitled “Bluff Dale ISD Outreach.”

Figure 100: Students Enjoying the Art Station, Decorating Parachutes

Figure 99: Students Learning at the Recovery Station at Dublin

Middle School

https://www.facebook.com/TarletonAeronauticalTeam/photos_albums


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VI) Project Plan

February 23, 2013

Tarleton Aeronautical Team was a guest speaker at the 2013 Tarleton Regional Science Olympiad. Students in attendance were highly involved in science curriculum. The team gave a STEM presentation and why it is important to students. Next, teams discussed the NASA USLI competition, the STEM aspects of this competition with physics, math computer science, engineering, etc. Other fields were discussed with their importance, such as English for writing documents, technical editing, and communication among team members as well as will the NASA review panel. The team then mingled with students and discussed personal interests in STEM, science, and the Tarleton Olympiad. No formal evaluation was conducted for this event; however, the meet and greet with the students proved to be informative and inspired interest among the students. Several schools inquired about middle school and high school eligibility to participate in the NASA SLPs.


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VII) Conclusion The Flight Readiness Review represents the final design that has been developed from project initiation, and has been demonstrated to meet all requirements set forth by the Statement of Work. This was verified by the successful full scale demonstration flight. The team confides in the overall design and is ready to launch on the official day, where flight performance will truly determine the integrity of the final project design. Up to this point, the project has been both challenging and rewarding. The mission complexity and project constraints have required creative, legitimate solutions to be 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 for both safety and reliability. All components of the vehicle aim to achieve the goal of delivering the SMD payload to one mile above ground level. The SMD payload criterion calls for an atmospheric data gathering instrument that meets all requirements as stated in the SOW. Orchestrating sensor control and data measurement successfully is the fundamental challenge for which solutions must be realized. This document represents the team‟s best effort in the design, test, and creation of payload that will meet the SMD requirements. 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. Ultimately, flying the vehicle and payload on launch day will illustrate the quality of the final design. With the results of the CDR, the team is eager to progress to the final stages of the design life cycle. It has been a truly rewarding experience to be involved in such a technical and complex project, and certainly provides great experience in working through the life cycle of a real-world engineering project.

flight readiness review - tarleton.eduflight readiness review i note to reader: to facilitate the...

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