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Enclosed: Flight Readiness Review

Submitted by:

2016 – 2017 Rocket Team Project Lead: David Eilken

Submission Date:

March 03, 2017

Payload: Fragile Material Protection

Mentor: Dr. David Unger, NAR 89083SR Level 2

Submitted to:

NASA Student Launch Initiative Program Officials

Faculty of the UE Mechanical Engineering Program

University of Evansville

College of Engineering and Computer Science

1800 Lincoln Avenue; Evansville, Indiana 47722

University of Evansville Student Launch

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Table of Contents

Table of Contents ...................................................................................................................... ii

List of Figures ........................................................................................................................... v

List of Tables ......................................................................................................................... viii

Nomenclature .......................................................................................................................... xii

FRR Summary .......................................................................................................................... 1

Design Updates from Proposal ................................................................................................. 2

Changes Made to Vehicle Criteria ........................................................................................ 2

Changes Made to Payload Criteria ........................................................................................ 2

Changes Made to Project Plan .............................................................................................. 3

Vehicle Criteria ......................................................................................................................... 4

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

Recovery.............................................................................................................................. 30

Mission Performance Predictions ....................................................................................... 34

Mission Performance Criteria ......................................................................................... 34

Flight Simulations and Altitude Predictions ................................................................... 34

Validity Assessment........................................................................................................ 41

Actual Stability Margin................................................................................................... 44

Kinetic Energy ................................................................................................................ 46

Drift ................................................................................................................................. 46

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Full Scale Flight .................................................................................................................. 47

Launch Day Conditions .................................................................................................. 47

Flight Analysis ................................................................................................................ 48

Flight Results .................................................................................................................. 54

Payload Criteria ...................................................................................................................... 56

Safety ...................................................................................................................................... 59

Personnel Hazard Analysis.................................................................................................. 60

Failure Modes and Effects Analysis.................................................................................... 71

Environmental Considerations ............................................................................................ 97

General Risk Assessment .................................................................................................. 102

Launch Operations Procedures ............................................................................................. 105

Parts Checklist ................................................................................................................... 105

Final Assembly Checklist ................................................................................................. 111

Motor Preparation ............................................................................................................. 117

Recovery Preparation ........................................................................................................ 118

Setup on Launch Pad ......................................................................................................... 120

Ignitor Installation ............................................................................................................. 121

Launch Procedures ............................................................................................................ 123

Troubleshooting ................................................................................................................ 125

Post-Flight Inspection ....................................................................................................... 127

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Project Plan ........................................................................................................................... 129

Testing ............................................................................................................................... 129

Altimeter ....................................................................................................................... 130

MTS (Bulkhead) ........................................................................................................... 131

Ejection Testing ............................................................................................................ 134

Parachute Deployment Force Testing ........................................................................... 136

Wind Tunnel Testing .................................................................................................... 137

Scale Model Testing ..................................................................................................... 148

Payload Testing ............................................................................................................. 148

Full Scale Testing ......................................................................................................... 156

Requirements Compliance ................................................................................................ 156

Budgeting and Timeline .................................................................................................... 174

Budget ........................................................................................................................... 174

Schedule ........................................................................................................................ 177

References ............................................................................................................................. 179

Appendix A – Machine Prints............................................................................................... 181

Appendix B – OpenRocket Simulation................................................................................. 200

Appendix C – Best Fit Curve ................................................................................................ 203

Appendix D – OpenRocket Simulation ................................................................................ 205

Appendix E – Payload Part Specification ............................................................................. 214

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Appendix F – Line Item Budget ........................................................................................... 218

Appendix G – Task Breakdown ............................................................................................ 219

Appendix H – Electrical Diagrams ....................................................................................... 223

Appendix I – Payload Accelerometer Graphs ...................................................................... 224

Appendix J – Wind Tunnel Uncertainty ............................................................................... 228

Appendix K – MTS Tensile Test Procedure ......................................................................... 230

List of Figures

Figure 1 - Recovery electronics wiring diagram ....................................................................... 6

Figure 2 - Altimeter Wiring ...................................................................................................... 7

Figure 3 - Body Tube .............................................................................................................. 12

Figure 4 - Fin Stops................................................................................................................. 13

Figure 5 - Painted Fins ............................................................................................................ 14

Figure 6 - Epoxy Nuts ............................................................................................................. 14

Figure 7 - Finished Nosecone ................................................................................................. 15

Figure 8 - Epoxy Location for the Centering Rings ............................................................... 16

Figure 9 - Complete coupling tube bulkhead assembly .......................................................... 18

Figure 10 - Coupling tube with permanent bulkhead and all thread rods ............................... 19

Figure 11 - Electronics sled assembly .................................................................................... 20

Figure 12 - Aft recovery mounting point ................................................................................ 21

Figure 13 - Cylinder 2 pin holes 3 inch spacing ..................................................................... 22

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Figure 14 - Bulkhead Cylinder 2 ............................................................................................ 22

Figure 15 - Rough finish on bulkhead .................................................................................... 23

Figure 16 - Wire rope isolator pin and aluminum square assembly ....................................... 24

Figure 17- Final Assembly...................................................................................................... 24

Figure 18 - Altimeter mounting assembly .............................................................................. 26

Figure 19 - Mounted O Ring ................................................................................................... 27

Figure 20 - Battery Holder ...................................................................................................... 28

Figure 21 - Mounting Pins ...................................................................................................... 29

Figure 22 - Altimeter Mounting Assembly ............................................................................. 30

Figure 23 - Block diagram of recovery system ....................................................................... 32

Figure 24 - Full-Scale Simulation ........................................................................................... 36

Figure 25 - Flight Simulation Input Data ................................................................................ 36

Figure 26 - Simulated Flight Configurations .......................................................................... 37

Figure 27 - Anticipated Motor Thrust Curve from OpenRocket ............................................ 39

Figure 28 - Flight 1 Actual vs OpenRocket Data ................................................................... 42

Figure 29 - Flight 2 Actual vs OpenRocket Data ................................................................... 43

Figure 30 - Flight 3 Actual vs OpenRocket Data ................................................................... 43

Figure 31 - Actual Cp and Cg locations.................................................................................. 45

Figure 32 - Actual Altitude vs OpenRocket Altitude ............................................................. 49

Figure 33 - Actual Data vs Regression ................................................................................... 50

Figure 34 - OpenRocket Data vs Regression ........................... Error! Bookmark not defined.

Figure 35 – Predicted Coefficient of Drag During Flight ....................................................... 53

Figure 36 - Final Design Assembly (new bolt and washer mounting) ................................... 56

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Figure 37 - Exploded view of payload assembly (annotation following) ............................... 57

Figure 38 - Exploded view base spring attachment ................................................................ 57

Figure 39 - CR1-400 wire rope isolator pin and plate assembly ............................................ 58

Figure 40 - Fracture mechanics clevis grip attached to U bolts Bulkhead assembly for MTS

testing .......................................................................................................................................... 132

Figure 41 - The Assembly Mounted into the MTS Machine ................................................ 133

Figure 42 – Variable Frequency Drive ................................................................................. 139

Figure 43- Strain Gage (From Vishay website) .................................................................... 140

Figure 44 - Strain Indicator ................................................................................................... 140

Figure 45 - Air Fan ............................................................................................................... 140

Figure 46 - Wind Tunnel....................................................................................................... 141

Figure 47 - Example of wiring strain gage to strain indicator .............................................. 141

Figure 48 - Wiring Diagram (strain gage to strain indicator) ............................................... 142

Figure 49 - Pareto Chart ........................................................................................................ 147

Figure 50 - Accelerometer Data Full-scale Flight 1 ............................................................. 155

Figure 51 - Sectional Budget Amounts ................................................................................. 176

Figure 52 - Gantt Chart ......................................................................................................... 178

Figure 53 – Aft Body Tube Drawing .................................................................................... 182

Figure 54 - Bow Body Tube Drawing .................................................................................. 183

Figure 55 - Fin Drawing ....................................................................................................... 184

Figure 56 - Motor Drawing ................................................................................................... 185

Figure 57 - Nosecone Drawing ............................................................................................. 186

Figure 58 - Launch Vehicle Drawing ................................................................................... 187

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Figure 59 – Recovery bulkhead drawing .............................................................................. 188

Figure 60 - Payload Main bulkhead residing in Cylinder 2 .................................................. 189

Figure 61 - Payload assembly general dimensions ............................................................... 189

Figure 62 - Recovery attachment bulkhead and hardware .................................................... 190

Figure 63 - Altimeter Mounting Plate Piece 1 ...................................................................... 191

Figure 64 - Altimeter Mounting Plate Vertical 1 .................................................................. 192

Figure 65 - Metal O-Ring ..................................................................................................... 193

Figure 66 – Propulsion Section ............................................................................................. 194

Figure 67 –Inner Tube .......................................................................................................... 195

Figure 68 - Centering Ring ................................................................................................... 196

Figure 69 - Thrust Plate ........................................................................................................ 197

Figure 70 - Inner Cylinder .................................................................................................... 198

Figure 71 - Payload Coupler ................................................................................................. 199

Figure 72 – 90 Degree Cotton Fill Large Bulb ..................................................................... 224

Figure 73 – 90 Degree Paper Fill Large Bulb ....................................................................... 225

Figure 74 – 90 Degrees Packing Peanuts Large Bulb........................................................... 225

Figure 75 – 90 Degrees Large Bulb Only ............................................................................. 226

Figure 76 – 90 Degrees DogBrag Fill Large Bulb ................................................................ 227

Figure 77 – 90 Degrees Base Value ...................................................................................... 227

List of Tables

Table 1 - Vehicle Specifications ............................................................................................... 4

Table 2 - System Level Functional Requirements .................................................................... 9

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Table 3 - Simulation Summary | Different Launch Configurations ....................................... 35

Table 4 - Rail Exit Velocity on Different Flights ................................................................... 39

Table 5 - Mach Number on Different Flights ......................................................................... 40

Table 6 – Impact of Wind Speed on Altitude ......................................................................... 41

Table 7 - Actual Stabilities ..................................................................................................... 45

Table 8 - Predicted kinetic energy of launch vehicle sections ................................................ 46

Table 9 - Predicted drift distance for selected wind speeds .................................................... 47

Table 10 - Actual Flight vs Predicted Flights Summary......................................................... 48

Table 11 - Definitions for Hazard and Failure Mode Analyses .............................................. 60

Table 12 - Personnel Hazard Analysis - Epoxy ...................................................................... 62

Table 13 - Personnel Hazard Analysis - Launch Operations/Post-Launch Inspection ........... 63

Table 14 - Personnel Hazard Analysis - Testing .................................................................... 66

Table 15 - Personnel Hazard Analysis - Fabrication .............................................................. 67

Table 16 - Personnel Hazard Analysis - Education Engagement Outreach Events ................ 70

Table 17 - Failure Modes and Effects Analysis - Design/Fabrication .................................... 72

Table 18 - Failure Modes and Effects Analysis - Payload...................................................... 76

Table 19 - Failure Modes and Effects Analysis - Payload Integration ................................... 80

Table 20 - Failure Modes and Effects Analysis - Recovery System ...................................... 83

Table 21 - Failure Modes and Effects Analysis - Testing ...................................................... 87

Table 22 - Failure Modes and Effects Analysis - Launch Support Equipment ...................... 90

Table 23 - Failure Modes and Effects Analysis - Launch Operations .................................... 94

Table 24 - Environmental Consideration Hazard Analysis .................................................... 97

Table 25 - General Risk Assessment .................................................................................... 102

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Table 26 - Parts Checklist - Propulsion ................................................................................ 105

Table 27 - Parts Checklist - Aerodynamics .......................................................................... 106

Table 28 - Parts Checklist - Main Payload ........................................................................... 107

Table 29 - Parts Checklist – Electronics Payload/Avionics Bay .......................................... 108

Table 30 - Parts Checklist - Recovery .................................................................................. 109

Table 31 - Parts Checklist - Safety and Education ............................................................... 110

Table 32 - Parts Checklist - Miscellaneous........................................................................... 110

Table 33 - Final Assembly Checklist - General Set Up ........................................................ 111

Table 34 - Final Assembly Checklist - Comprehensive Structural Inspection ..................... 112

Table 35 - Final Assembly Checklist - Electronics .............................................................. 112

Table 36 - Final Assembly Checklist - Payload.................................................................... 113

Table 37 - Final Assembly Checklist - Recovery System .................................................... 113

Table 38 - Final Assembly Checklist - Motor/Ejection System Preparation ........................ 115

Table 39 - Final Assembly Checklist - Secure Attachment Inspection ................................ 116

Table 40 - Final Assembly Checklist - Launch Pad/Pre-Launch Inspection ........................ 116

Table 41 - Motor Preparation Checklist ................................................................................ 117

Table 42 - Recovery Preparation Checklist .......................................................................... 118

Table 43 - Launch Pad Configuration Checklist................................................................... 120

Table 44 - Ignitor Installation Checklist ............................................................................... 121

Table 45 - Launch Procedures Checklist .............................................................................. 123

Table 46 - Troubleshooting - Cracking in Main Body Tube or Subsection ......................... 125

Table 47 - Troubleshooting - Insecure Fit Between Adjoining Subsections ........................ 125

Table 48 - Troubleshooting - Unresponsive or Malfunctioning Electronics ........................ 126

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Table 49 - Troubleshooting - Insecure Connection Between Launch Rail and Launch Pad 126

Table 50 - Post-Flight Inspection Checklist.......................................................................... 127

Table 51 - Test Results ......................................................................................................... 129

Table 52 - MTS Test Results ................................................................................................ 134

Table 53 - Results of ejection testing .................................................................................... 136

Table 54 - Maximum force on launch vehicle during descent .............................................. 137

Table 55 - Testing Apparatus Components .......................................................................... 138

Table 56 - Inputs for Uncertainty analysis ............................................................................ 146

Table 57 - Spring Constant Test Values ............................................................................... 149

Table 58 - Charpy Impact Acceleration Test Data ............................................................... 151

Table 59 - Fragile Material Sample Testing ......................................................................... 153

Table 60 - Full Scale Flight Results ...................................................................................... 156

Table 61 - NASA Requirement Compliance ........................................................................ 157

Table 62 - Team Requirement Compliance .......................................................................... 169

Table 63 - Sources of Funding .............................................................................................. 174

Table 64 - Sectional Budget Breakdown .............................................................................. 175

Table 65 - Critical Dates ....................................................................................................... 177

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Nomenclature

𝐴𝑃 Ammonium Perchlorate Composite

𝐵𝐻 Bulkhead

𝐶𝑝 Specific Heat with constant pressure [kJ/kmol-K]

𝐶𝑅 Centering Rings

𝐸 modulus of elasticity [psi]

𝐹 Force [lbf]

𝑓 operational frequency [Hz]

𝑓𝑛 natural frequency [Hz]

𝐹. 𝑂. 𝑆 Factor of Safety

ℎ enthalpy [kJ/kmol] Combustion Analysis Section

ℎ thickness [in]

ℎ𝑓𝑜 Enthalpy of Formation [kJ/kmol]

𝐼 second moment of inertia [𝑖𝑛4]

𝑘′ stiffness [psi]

KE Kinetic Energy

𝐿 length [in]

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𝑀 moment [lbf-in]

𝑚 mass [lbm, kg] kg will be specified in the equation otherwise it is lbm

𝑛 moles [kmol]

𝑃 pressure [kPa]

𝑄 heat [kJ]

𝑅 radius [in]

𝑅 Gas Constant [=8.314 kJ/kmol-K] Combustion Analysis Section

𝑟 frequency ratio

𝑆𝐴 surface area [𝑖𝑛2]

𝑇 temperature [K]

𝑡 time [s]

𝑉 volume [𝑖𝑛3]

𝑣 velocity of the rocket at burnout [m/s]

𝑣𝑓 ground impact velocity [ft/s]

𝑣𝑑 descent rate [ft/s]

𝜔 circular natural frequency [rad/sec]

𝑤 work [kJ]

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FRR Summary

Project ACE will field a 111.75” long, 32.6-pound carbon fiber and aluminum based rocket.

The leading tip of the rocket begins with a G-10 Fiberglass, 22”, ogive nosecone. Contained in a

pressure-equalizing compartment in the nosecone sits the official altimeter as well as a GPS

tracking system. Just aft of this compartment are four threaded rods for fastening ballast. A

fragile material protection system resides below the nosecone. The payload contains concentric

cylinders, connected by an array of springs and wire-rope isolators selected through extensive

mathematical modeling. The innermost cylinder, where the fragile material is to be contained,

features a variable position cap and fill material to ensure that the fragile material will be

contained under sufficient pressure regardless of volume. It is the team’s objective to produce a

successful payload that provides meaningful vibration and impulse reduction information.

Moving aft from the payload is the recovery system. This system features completely

redundant separation circuits. At apogee, a 24” drogue chute ejects, followed by a 96” main

chute at 750’. At the aft end of the rocket is the propulsion section. A 75-mm L-850W Aerotech

motor propels the rocket for just over four seconds to an altitude of one mile. A 12’ 1515

extruded aluminum launch rail has been selected to achieve an acceptable rail-exit velocity. The

motor is held in place via 6061-T6 Aluminum centering rings and thrust plate. All components

are housed in two carbon fiber body tubes. The fins, which adhere to the centering rings and

body tubes, are made out of G-10 Fiberglass and have a clipped delta design. Each system is

covered in much more depth in the “Vehicle Criteria” section of this report. For specific team

information, such as the mentor and mailing address, please see the cover page of this report. For

more “quick facts” on the rocket please reference the associated milestone review flysheet.

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Design Updates from Proposal

Changes Made to Vehicle Criteria

The nosecone shoulder was shortened from 5.25 inches to 3 inches. The change accounts for

spring oscillation from the main payload, which is located below the nosecone in the bow body

tube.

A 3

8 inch hole was drilled in the furthest aft centering ring. The hole allows the furthest aft rail

button to be tightened and loosened as needed by giving access to the interior threads of the rail

button. An aluminum nut holds the rail button to the airframe and the rail button assembly is now

removable if necessary through the aft centering ring.

It was observed that after sub-scale test flights, the quarter-inch quick links used to secure the

recovery harnesses to the launch vehicle body tubes had become mildly deformed, making it

irksome to tighten or loosen them. Consequently, larger quick links were initially selected for the

full-scale launch vehicle. The shift in mass associated with these larger quick links created a low

stability off the launch rod for the second full-scale test flight. In order to return stability to an

acceptable value, the original quarter-inch quick links were reemployed and will be used on all

launch configurations moving forward.

Changes Made to Payload Criteria

The main change the payload saw was the mounting of the base springs. Due to epoxy

failing during impact and welds weakening the integrity of the spring, a new design was

developed and used in testing. This design is spoken about in detail in the “Payload” section of

this report. One other decision that was made through the testing of the payload was the final

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choice of fill material for Cylinder 1 (the innermost cylinder). The selection is a combination of

shredded paper and cotton filling and the rationale behind the decision is described in detail in

the payload testing section of this report. During testing, epoxy continued to fail on mating

surfaces especially where the wire rope isolators were adhered to Cylinder’s 1and 2. To combat

this failure, pins were used and inserted into holes in both cylinders with epoxy to add further

strength. Thus, for the payload as a whole, conceptual changes were not made but small changes

to the mounting design were made.

Changes Made to Project Plan

A few changes were made to both the schedule and budget. Project ACE’s build phase

extended about one week longer than anticipated – mainly due to the redesign of payload

mounting. The redesign resulted in a multi-week delay in testing. Project ACE was also forced

to launch one week late due to weather. The team launched with “BluesRocks Rocketry” in

Elizabethtown, KY instead of Laünch Crüe. These scheduling changes are reflected in the

“Schedule” section of this report, but Project ACE is on schedule once again.

Slight alterations were made to the budget to accommodate sections with unforeseen

costs. In essence, funds for sections of the rocket that were under budget were allocated to

administrative and payload sections to cover overrun costs. Project ACE remains under budget,

but more detail on the allocation of funds can be found in the “Budget” section of the report.

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Vehicle Criteria

Design and Construction of Vehicle

Design Features

Structural Elements

Vehicle Overview

The vehicle specifications can be seen in Table 1. The overall length of the rocket is 111.75

inches with a diameter of 5.5 inches.

Table 1 - Vehicle Specifications

Component Dimension Material

Bow Body Tube 48 inches Carbon Fiber

Aft Body Tube 41 inches Carbon Fiber

Nosecone 21.75 inches G10 Fiberglass

Bulkhead/Centering Ring ¼ inch Aluminum

Coupler 12 inches G10 Fiberglass

Body Tubes

The body tubes provide the structural rigidity necessary for housing the internal components

as well as undergoing flight/recovery stresses. These tubes also account for the bulk of the mass

of the airframe and provide a large surface area for airflow while in flight. To guarantee a

successful flight, all these factors must be accounted for in the material selection of the body

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tubes. Carbon fiber was selected to provide a lightweight frame for the launch vehicle (0.658

oz/in3) while also providing a higher tensile strength than that of fiberglass or BlueTube.

Nosecone

The nosecone must withstand the forces of in-flight airflow and vehicle recovery, however,

both of these forces are minimal and do not require the increased strength provided by carbon

fiber. The nosecone is also smaller than the body tubes and provides less of a weight reduction

from carbon fiber to fiberglass. Lastly, the official scoring altimeter of Project ACE is housed in

the nosecone. The altimeter is specified to not be housed in carbon fiber for transmission

purposes. For these reasons, it was chosen to use fiberglass instead of carbon fiber.

Coupling Tube

The coupling tube serves as the joint between body tubes and the housing for the recovery

electronics. This coupler separates from the remainder of the launch vehicle during the recovery

process. Aluminum caps seal the space on either side of the coupler, and threaded aluminum rods

connect the aluminum caps. The caps and rods bear the stresses of the recovery process. For this

reason, an inexpensive material was able to be chosen for the coupler. BlueTube was the

material chosen because it is readily available from many manufacturers at a low cost.

Bulkheads/Centering Rings

The bulkheads and centering rings provide additional structural integrity for the launch

vehicle. They also serve as possible mounting points for vehicle components such as the motor

retention system or shock cords. Lastly, they are used to separate the recovery section from the

payload and the propulsion section. Aluminum was chosen for the bulkheads for its high tensile

strength (300 MPa) to ensure the success of the crucial functions these components perform.

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Electrical Elements

Recovery electronics are connected using 20-gauge wire. Screw terminals are used to make

electrical connections with the terminal blocks and recovery altimeters. Connections to the rotary

arming switches are soldered. 4-pin Molex connectors are soldered in series with the signal wires

from each altimeter to allow the electronics sled to be removed from the coupling tube. To allow

for easier management, all wire pairs were twisted neatly and some wiring was secured to the

sled using metal retainers. A complete wiring diagram of the recovery electronics is shown in

Figure 1.

Figure 1 - Recovery electronics wiring diagram

The scoring altimeter has two connections. The first, which runs to the battery, connects

into the socket at the bottom of the altimeter. On the other end of the wire, it is soldered into the

battery. The second connection is the switch to turn the altimeter on or off. The toggle switch is

soldered to two leads which are locked into place on the altimeter by a terminal block. Both the

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attached toggle switch and the battery connection can be seen on Figure 2 in its respective

socket.

Figure 2 - Altimeter Wiring

Drawing and Schematics

A full listing of dimensioned drawings can be seen in Appendix A.

Flight Reliability

Mission Success Criteria

Listed below are the mission success criteria determined by the Project ACE team.

1. Aerodynamics

a. The airframe, nose cone, and fins remain intact for the duration of the flight.

b. The airframe, nose cone, and fins are reusable for any following flights.

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c. The airframe and nose cone will protect all internal components from damage

from external sources.

2. Propulsion

a. The vehicle will attain an apogee between 5,125 feet and 5,375 feet.

b. The vehicle will remain below Mach 1.

c. The motor mount will withstand propulsion forces and remain reusable for

any following flights.

3. Recovery

a. The drogue parachute and main parachute are ejected at apogee and 750 feet,

respectively.

b. The drogue parachute and main parachute inflate successfully following

ejection.

c. The maximum kinetic energy of any independent section of the rocket is less

than 75 ft-lbf at landing.

4. Electronic Payload

a. The data sent from the electronic payload is received remotely during and

after the vehicle’s flight.

b. The electronic payload withstands flight forces and remains reusable for any

following flights.

c. The electronic payload accurately determines the apogee of the rocket.

5. Main Payload

a. The fragile object(s) remain undamaged.

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b. The force acting on the payload is reduced by 50% for each of the areas of

interest: (thrust curve, parachute deployment, and landing.)

c. The acceleration acting on the payload is reduced by 50% for each of the areas

of interest: (thrust curve, parachute deployment, and landing.)

Flight Reliability Confidence

The system level functional requirements can be seen in Table 2 where the severity and

likelihood of failure in each mission success criteria and the action performed to mitigate these

failures are described.

Table 2 - System Level Functional Requirements

Section Success Criteria Explanation Severity

of Failure

Likelihood

of Failure

Aer

odynam

ics

The airframe, nose cone,

and fins should remain

intact for the duration of

the flight.

A failure of the airframe during

flight could cause a complete

failure in the launch vehicle’s

flight ability. However, the use

of carbon fiber mitigates this risk

to a very low likelihood.

Significant Low

The airframe, nose cone,

and fins should be

reusable for any

following flights.

Reusability of parts is not

detrimental to the project; new

parts can be purchased. The most

likely part to fail is a fin upon

recovery, thus warranting a

medium likelihood of failure.

Minor Medium

The airframe and nose

cone should protect all

internal components

from damage from

external sources.

Damage to internal components

can be detrimental to the launch

vehicle’s ability to deploy the

recovery system. However, a

carbon fiber airframe mitigates

this risk to a very low likelihood.

Major Low

Pro

puls

ion

The vehicle should

attain an apogee

between 5,125 feet and

5,375 feet.

Motor variations and launch day

conditions both contribute to

apogee variations from the full-

scale test. However, the team’s

allocated window should

account for these.

Medium Low

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Section Success Criteria Explanation Severity

of Failure

Likelihood

of Failure

The vehicle should

remain below Mach 1.

This is a requirement from

NASA. The vehicle has not been

designed to withstand transonic

forces. The anticipated Mach

number is 0.56.

Significant Low

The motor mount should

withstand propulsion

forces and remain

reusable for any

following flights.

Motor mount/retention failure

could cause a poor flight, no

flight, or safety hazard. This is

mitigated by using aluminum

and high strength epoxy.

Major Low

Rec

over

y

The drogue parachute is

successfully deployed at

apogee.

If the drogue parachute does not

deploy at apogee, the main

parachute will deploy at high

velocity. This could result in

damage to the parachute or

airframe.

Major Low

The main parachute is

successfully deployed at

750 feet.

If the main parachute does not

deploy, the launch vehicle will

descend under only the drogue

parachute. This would result in

excessive ground impact speed.

Major Low

The drogue parachute

and main parachute

inflate successfully

following ejection.

A partially-inflated parachute is

much less effective at slowing

the launch vehicle during its

descent. This could result in

excessive ground impact speed.

Major Low

The maximum kinetic

energy of any

independent section of

the rocket is less than 75

ft-lbf at landing.

Excessive kinetic energy on

landing could result in damage

to the fragile payload or

airframe.

Major Low

Ele

ctro

nic

Pay

load

The data sent from the

electronic payload

should be able to be

received remotely

during and after the

vehicle’s flight.

If the data is not received

remotely after the flight, the

team will not be scored.

Low Low

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Section Success Criteria Explanation Severity

of Failure

Likelihood

of Failure

The electronic payload

should withstand flight

forces and remain

reusable for any

following flights.

The altimeter notwithstanding

the forces will prevent the

altimeter from being reused, and

the team cannot be scored.

Low Low

The electronic payload

should accurately

determine the apogee of

the rocket.

If apogee is not detected

accurately, it will affect the score

of the team.

Low

Low

Mai

n P

aylo

ad

The fragile object(s)

should remain

undamaged.

The force felt by the

payload should be

reduced by 50% for

each of the areas of

interest: takeoff (thrust

curve, parachute

deployment, and

landing.)

To properly reduce the risk of

damage to any and all unknown

fragile material, the desired

reduction of force felt by the

payload should be reduced by a

minimum of 50 percent for the

most extreme forces exerted

throughout flight.

Major Medium

The Acceleration felt by

the payload should be

reduced by 35% for

each of the areas of

interest: (thrust curve,

parachute deployment,

and landing.)

To reduce the force at the

maximum and minimum points

of spring displacement, total

acceleration of the payload

should be reduced by a

minimum of 35 percent.

Significant Low

Construction Process

Body Tubes / Nosecone

One of the first steps that was taken in the construction of the body tubes was to cut down

the bow body tube to forty-one inches in length, using a horizontal band saw. This can be seen in

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Figure 3. The body tube was then filed smooth using a file. The next step was drilling three 1

8

inch diameter holes in the bow body tube using a CNC. These holes were used to bolt the

nosecone into the bow body tube. Three holes were drilled for the rail buttons on the side of the

aft body tube using a CNC.

Figure 3 - Body Tube

The locations of the rail buttons were determined using the center of gravity and center of

pressure. The rail buttons are attached to the body tube using a nut and bolt. Two of the three rail

buttons are accessible, and can be removed from the rocket. The first rail button was fastened

onto the aft body tube in the recovery section, the second rail button is fixed on the exterior of

the aft body tube between two bulk heads, and the third is accessible through a hole that was

drilled in the lower bulk head. The second rail button was not only bolted together, but also

epoxied in place to ensure that it would not move or break free. The aft body tube was slotted

using a CNC. These slots allowed the fin tabs to be inserted into the body tube.

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The fins were made from two, 2’x1’x1/8” sheets of G10 fiberglass. These were hand cut to

the desired fin dimensions and were beveled using a freestanding horizontal belt sander. The

dimensions of the fins are: 5.5 inches tall, the tip chord is 5.8 inches long, the root chord is 7.5

inches long, the fin tab has a height of 1.2 inches tall, and the fin tab length is five inches. 3

8 inch

thick ABS Plastic fin stops, pictured in Figure 4, were manufactured using a 3-D printer. The fin

stops are inserted into the body tube flush with the centering ring, and epoxied on all contact

surfaces to ensure a solid fit. These fin stops fit between each of the already epoxied fins, and

provide extra internal support for the fins. Then the fins were painted for aesthetics, seen in

Figure 5.

Figure 4 - Fin Stops

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Figure 5 - Painted Fins

The shoulder of the nosecone was reduced to three inches to allow the main payload to

oscillate. The nosecone was attached to the body tube using three epoxy nuts, and bolts. An

example of an epoxy nut can be seen in Figure 6. These epoxy nuts were epoxied using

RocketPoxy to the inner diameter of the body tube, concentric with the bolt hole. The holes in

the nosecone were lined up with the holes in the body tube before the bolts were tightened to

form a snug fit. The nosecone was also painted for aesthetics, as seen in Figure 7.

Figure 6 - Epoxy Nuts

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Figure 7 - Finished Nosecone

Propulsion

The propulsion section was constructed over a period of five days to ensure the epoxy

was completely set before moving to the next section of the construction process. The first part

of the process was milling out the center of the engine block (bulkhead) to be certain that the

inner tube was in the middle of the plate. After the milling was completed, the blue tube cut to a

length of 21.75 inches was epoxied to the milled portion of the engine block and allowed to cure

overnight.

The next day, the engine block and inner tube assembly were epoxied into the body of the

rocket 21 in from aft end of the rocket. The bulkhead was epoxied on the bow and aft side for a

secure bond. Once the epoxy was applied, a loose centering ring was placed at the aft end of the

body tube to make sure the bulkhead and epoxy were set completely in-line with the body tube.

Once the bulkhead epoxy dried, the first centering ring was epoxied to the body tube 19 in from

aft end of the rocket. The centering ring was epoxied on both the outer and inner edge around the

body tube and the inner tube for a secure bond. Again, a loose centering ring was applied to

ensure the centering right would not set-up at an angle. Figure 8 shows the location where the

epoxy is applied to the centering rings.

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Figure 8 - Epoxy Location for the Centering Rings

Once the first centering ring was set, the centering ring that sits in front of the fin tables

was inserted. A fishing wire was tied around the centering ring so it could be pulled back to the

front of the tabs for adjustability. When the fins were placed into the body tube, the centering

ring was pulled aft to sit against the front of the fin tabs. With the centering ring in the correct

position, the fins were removed to epoxy the circumference of the outer and inner edges of the

centering ring. Once the epoxy was applied, a loose centering ring was inserted over the inner

tube to ensure the entire assembly would not fall at an angle with the epoxy setting.

With the centering ring dry, epoxy was applied to the area where the fins would sit

against the centering ring. A more detailed description on how the fins and fin stops were

assembled is included in the aerodynamic construction section. When all fins were epoxied in

place, the last centering ring was inserted onto the inner tube resting against the back of the fin

Outer Area Epoxy Location

Inner Area Epoxy Location

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tabs. The centering ring was epoxied in place. Finally, epoxy was applied to the inner surface of

the retention system and then placed onto the inner tube. Then epoxy was applied to the outside

of the retention system making a fillet between the blue tube and the retention system.

Recovery

The construction of the full-scale recovery system began with the assembly of the coupling

tube. A stock 12-inch blue tube coupler of 5.48-inch OD was selected to house the electronics.

First, a 1-inch ring of 5.5-inch OD blue tube was epoxied around the middle of the coupling tube.

This ring provides a smooth, continuous surface as air passes from one body tube to the next. It

also locates the coupling tube vertically within the body tubes and allows access to the recovery

electronics through pressure sampling holes. Four pressure sampling holes of 0.286-inch

diameter (as specified by the PerfectFlite Stratologger CF manual) were drilled through the ring

and coupling tube, spaced equally around the circumference of the tube. Finally, two smaller

rings of blue tube were epoxied to the inside of each end of the coupling tube, leaving a 0.25-

inch shoulder to locate the bulkheads.

Two 0.25-inch thick aluminum bulkheads of 5.175-inch OD were machined using a CNC

mill. Holes of 0.25-inch and 0.3125-inch diameter were drilled on perpendicular axes to

accommodate all thread rods and U-bolts, respectively. Each bulkhead received a steel U-bolt

secured with hex nuts, flat washers, a steel backing plate, and lock washers. Epoxy was applied

around the washers and nuts after assembly to ensure the bulkhead was airtight. One bulkhead,

hereafter referred to as the “permanent bulkhead”, received two 14-inch steel all thread rods

secured with hex nuts, flat washers, and lock washers. The all thread rods were first located such

that they spanned the entire length of the coupling tube with equal lengths protruding from each

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bulkhead when fully assembled. They were then epoxied to the permanent bulkhead with all

mounting hardware previously described. Two ejection charge wells made from 1-inch OD

aluminum tubing were then epoxied to the outside of each bulkhead. A completed bulkhead is

shown in Figure 9.

Figure 9 - Complete coupling tube bulkhead assembly

After the all thread rods had been permanently fixed, the construction of the electronics sled

could begin. First, a brass tube of 0.25-inch ID was slid over each all thread rod. The brass tubes

each received a thin bead of epoxy along their lengths before being pressed against a sheet of

balsa wood. Once dry, the tubes were correctly located to match the all thread rods, and

additional epoxy was applied to secure the balsa wood sled to the brass tubes. Mounting holes

were then drilled to accommodate the altimeters and batteries. The altimeters were secured using

#4 bolts and nuts, while the batteries were secured using plastic zip-ties. Next, each altimeter

U-BOLT EJECTION WELL

MOLEX CONNECTOR

TERMINAL BLOCK

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arming switch was mounted through a hole in a small piece of balsa wood. This assembly was

epoxied to the electronics sled such that the switched faced out radially, in line with opposite

pressure sampling holes.

Next, the permanent bulkhead was epoxied into the coupling tube at an orientation that

aligned the arming switches with the pressure sampling holes, being careful to create an airtight

connection around the circumference of the bulkhead. A bead of silicone rubber was applied

around the shoulder where the other bulkhead, hereafter referred to as the “removable bulkhead”,

was to rest. This ensured an airtight seal around the removable component. The coupling tube

assembly with permanent bulkhead and all thread rods is shown in Figure 10.

Figure 10 - Coupling tube with permanent bulkhead and all thread rods

With the electronics sled removed from the coupling tube, wires were soldered to each

terminal of the arming switches and connected to the dedicated switch leads of each altimeter.

MOLEX CONNECTOR

ALL THREAD ROD

COUPLING TUBE

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Connections were made between each battery and the dedicated power leads of the

corresponding altimeter. To allow for easy replacement of spent igniters, terminal blocks were

epoxied to the outside of each coupling tube bulkhead. A 0.125-inch hole was then drilled in

each bulkhead to allow the passage of wires from the interior of the tube to the terminal blocks.

Two pairs of wires (one for each recovery event’s redundant igniters) were connected to each

terminal block and fed through the bulkheads before being epoxied to create an airtight seal, as

shown previously in Figure 9. The four wires concerned with main parachute deployment (from

the aft-most coupling tube bulkhead) were connected in pairs to the primary and backup

altimeter’s MAIN leads, and the four wires concerned with drogue parachute deployment (from

the bow-most coupling tube bulkhead) were connected in pairs to the primary and backup

altimeter’s DROG leads. In order to allow for the easy removal of the electronics sled between

flights, these connections were made impermanent using 4-pin Molex connectors soldered

between the altimeter leads and the terminal blocks. These connectors can be seen in Figure 9

and Figure 10, while the entire electronics sled assembly is shown in Figure 11.

Figure 11 - Electronics sled assembly

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After construction of the coupling tube and its associated systems was complete, the two

permanent recovery mounting points in the body tubes were created. Each mount was created by

epoxying a steel U-bolt through an aluminum bulkhead of 0.25-inch thickness and 5.35-inch OD

using the same hardware described for the coupling tube bulkheads. Each recovery mount was

secured by first applying a small amount of epoxy around the inside of the body tube where the

bulkhead would be located. After pressing each bulkhead into its final location, a small fillet of

epoxy was applied around its circumference, followed by a larger fillet once the first had dried.

The aft-side recovery mounting point is shown in Figure 12 after being epoxied into place.

Figure 12 - Aft recovery mounting point

Main Payload

The assembly for the main payload began with cutting the 5.36” OD Blue Tube Coupler to

11 inches. Holes were then drilled into the coupler at a spacing of 3 inches apart starting 3 inches

from the bottom of Cylinder 2, as seen in Figure 13.

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Figure 13 - Cylinder 2 pin holes 3 inch spacing

The payload was also designed with two bulkheads, created with the CNC from 0.25-inch

aluminum. The first bulkhead was epoxied into Cylinder 2 and had fifteen 0.2-inch diameter

holes were milled and threaded for the bolts used to attach the base springs. Five 0.5-inch holes

were milled to center the springs and a 1-inch diameter hole was milled out of the center to

reduce the weight of any moving parts within the rocket. The final bulkhead can be seen in

Figure 14.

Figure 14 - Bulkhead Cylinder 2

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The second bulkhead mentioned was milled similarly, however instead of a 1.5-inch diameter

hole in the center for weight reduction, it had two 0.3-inch diameter holes for the bolts used to

connect to the recovery bulkhead located immediately below in the rocket.

Once the bulkheads were machined, the surface around the edge of the first was roughed up

with a file to increase surface area for the epoxy to hold to. This can be seen in Figure 15.

Figure 15 - Rough finish on bulkhead

It was then epoxied into Cylinder 2 and set to dry. Once dry, the first and second bulkheads

where mounted to the five base springs via the bolt and washer assembly spoken about in the

Payload section of this report. After the base springs were mounted, the wire rope isolators were

prepared, small 1-inch by 1-inch squares of 0.1-inch aluminum sheet metal were cut to be

epoxied to the 3D printed Cylinder 1 to prevent any failure in tension due to weaknesses in 3D

printed material. 0.3604-inch holes were drilled into the aluminum squares. The surface of each

square was roughed up with a file and then soaked in Acetone to clean prior to epoxying.

Cylinder 1 then had 0.3604” inch holes drilled into it at the same spacing as Cylinder 2 however

starting 2 inches from the base of Cylinder 1 to allow for a maximum oscillation of 1-inch within

Cylinder 2. Pins were cut using a hack saw from standard deck nails that happened to be the

correct size as the thru holes in the wire rope isolators. Each wire rope isolator was epoxied to a

2-mm long pin and then epoxied to the aluminum square, as shown in Figure 16.

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Figure 16 - Wire rope isolator pin and aluminum square assembly

After the epoxy cured on the wire rope isolators, the pins were inserted into the holes drilled

into Cylinder 1, the aluminum plate and pin assembly was epoxied to the plastic Cylinder. After

3 hours, Cylinder 1, now attached to all 12 wire rope isolators was inserted into Cylinder 2.

Epoxy was placed on all exposed pins and outer faces of the wire rope isolators, to adhere to the

inner diameter of Cylinder 2. The pins were inserted into the holes in Cylinder 2 and set to dry.

Prior to launch, two bolts were screwed in aft of the recovery bulkhead to secure the entire

payload assembly in the rocket. The final assembly can be seen in Figure 17.

Figure 17- Final Assembly

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Electronic Payload

The electronics payload consists of the altimeter, the battery, the mount, and the ballast

attach points. The mount consists of four components: the O-ring, base plate, vertical mounting

plate, and the battery holder. All four-mount components are machined from Aluminum 6061

and were milled on a 3-axis CNC mill. Once milled, the base plate and vertical mounting plate

were tig welded to form one assembly shown in Figure 18.

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Figure 18 - Altimeter mounting assembly

The O-ring serves as a permanent mounting point for the base plate. The base plate attaches

to the O-ring via 4 manually threaded holes. The O-ring was permanently fixed in the nosecone

using Rocketpoxy. The attached O-ring can be seen in Figure 19.

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Figure 19 - Mounted O Ring

Figure 20 shows the battery holder that was designed to attach to the assembly. The battery

holder was designed as a separate piece because the original assembly was to hold the battery

already. Mounting the battery under the altimeter would not allow the altimeter to accurately

measure altitude. Due to the inaccuracy, a new battery holder was designed to securely attach on

to the vertical mounting plate. Holes were made and tapped on the backside of the altimeter

mount to be the attach points.

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Figure 20 - Battery Holder

To allow adjustments based on the test flights actual performance compared to simulation,

the team needed the ability to add ballasts to the launch vehicle. Weights will be mounted to the

aft end of the base plate. Ballast mounts were designed in the base plate. Using the CNC mill,

four ballast holes were cut and then tapped. Mounting pins were coated in epoxy and screwed

into the holes. When the epoxy dried, remaining was the four mounting pins for ballasts to be

attached, as shown in Figure 21.

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Figure 21 - Mounting Pins

Clear plastic tubing was run from the bottom of the base plate to the base of the nosecone to

allow the nosecone compartment to properly pressurize during vehicle flight. The opposite end

of the tubing was attached to the outer wall of the nosecone shoulder using a PVC fitting. The

PVC fitting was attached using Rocketpoxy. The mounting assembly with the attached altimeter

is shown in Figure 22.

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Figure 22 - Altimeter Mounting Assembly

Recovery

The first recovery event is the deployment of a Fruity Chutes CFC-24 parachute at apogee.

This 24-inch-diameter ripstop nylon parachute will serve as the launch vehicle’s drogue

parachute, resulting in an initial descent velocity of 76.5 ft/s. The second recovery event is the

deployment of a Fruity Chutes IFC-96 parachute at 750’ above ground level. This 96-inch-

diameter ripstop nylon parachute serves as the launch vehicle’s main parachute, resulting in a

final descent velocity of 14.5 ft/s.

Two 35’ lengths of 1-inch tubular nylon are used as recovery harnesses to tether the three

independent sections of the launch vehicle together. To secure the harnesses to rocket, a loop is

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stitched at the end of each harness using Kevlar thread. An additional loop is stitched into each

harness 5’ from one end, which serves as an attachment point for each harness’s parachute.

Attachment hardware consists of 5/16” steel U-bolts secured to the bulkheads with lock washers

and steel backing plates to distribute loading during recovery events, as shown previously in

Figure 9.

The recovery events are controlled by two PerfectFlite StratoLogger CF altimeters. These

altimeters utilize a pressure transducer to determine the altitude of the launch vehicle. The

Stratologger CF is relatively simple, yet effective. It has the ability to fire two igniter signals: one

at apogee with an adjustable delay time and the other at a fixed altitude. This configuration is

ideal for dual-deployment. Using a software transfer kit, altitude and temperature data can be

obtained for up to 16 stored flights.

Separate 9-volt lithium-ion batteries are connected to the power terminals of each altimeter.

The Stratologger CF also has dedicated terminals for connecting a power switch. Using these

terminals, a rotary locking switch is connected and used to toggle power to each altimeter.

QuickBurst QBECS igniters are connected to the drogue and main output terminals of each

altimeter. These low-current igniters ensure reliable, complete ignition of the black powder

ejection charges.

Redundancy of the recovery system is achieved by utilizing two identical sets of components

with completely separate electrical circuits. In this way, if either circuit were to be shorted

accidentally or experience an altimeter malfunction, the other circuit would remain unaffected. In

addition to the redundant circuitry, each igniter is inserted into its own separate ejection charge

well with the appropriate amount of black powder. The result is two black powder explosions for

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each ejection event. To avoid over-pressurization of the parachute compartments, the ignition

signals from the backup circuit are delayed using the altimeter’s built-in software. The first

signal for the drogue parachute is fired at apogee and the backup signal is fired 2 seconds later.

The first signal for the main parachute is fired when the launch vehicle reaches an altitude of 750

feet and the backup signal is fired when it reaches 650 feet. In both scenarios, a successful

ignition at the primary signal results in the backup ejection charge exploding harmlessly into the

atmosphere. Conversely, if a main charge fails to ignite for any reason, the backup signal causes

ignition and subsequent parachute ejection due to pressurization of the parachute compartment.

A block diagram of the redundant recovery electrical systems is provided in Figure 23.

Figure 23 - Block diagram of recovery system

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The scoring altimeter uses the 470 MHz frequency bands to transmit the GPS and live

feed from the rocket. The GPS has an operational altitude limit of 50,000 meters. The scoring

altimeter requires 0.592 Watts to run during the flight. Project ACE recognized that interference

from the scoring altimeter to the recovery system is possible. To ensure that the interference

would not compromise the recovery system, all ejection tests were done with the scoring

altimeter on and near the body of the rocket. Keeping the scoring altimeter near the body would

allow any interference to affect the recovery system. In doing so, there was no noticeable

interference or change to the recovery system. In addition to the ejection testing, all altitude tests

with the drone had all three altimeters mounted in the same location. With all drone testing, there

was no noticeable interference with the recovery system.

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Mission Performance Predictions

Mission Performance Criteria

The main mission performance objective for the team is to reach an altitude between 5,125

and 5,375 feet. The goal gives the team a range of 250 feet which is an accomplishable goal for a

first-year team. Another goal for the launch transport a piece of fragile material and safely return

it back to Earth after reducing kinetic energy to less than 75 pound-force. The altitude range and

the fragile payload are but a few of the goals set forth by NASA and the team – see the

“Requirements Compliance” section for more. The goals were then measured through testing of

the full-scale rocket. The team used three altimeters, one located in the nosecone that can

measure acceleration, velocity, and altitude, and two located in the recovery bay measuring just

the altitude. An accelerometer was used to measure the fragile material payload bay force

reduction and the accelerometer in the nose cone is used to calculate the energy of the rocket as it

lands back on Earth.

Flight Simulations and Altitude Predictions

The full-scale rocket was tested three times with three different configurations. Both

ballast weight and quick link (in the recovery section) style was altered. The configurations can

be seen in Table 3. The different configurations were simulated in OpenRocket using the

conditions of the launch day to mimic actual conditions as closely as possible.

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Table 3 - Simulation Summary | Different Launch Configurations

Simulation Launch Day

Conditions Quick Links Ballast (lb)

Simulated

Apogee (ft)

1a – Baseline No Heavy 2.0 5,005

Flight 1 Yes Heavy 2.0 4,967

Flight 2 Yes Heavy 0.0 5,322

Flight 3 Yes Light 1.5 5,326

For a baseline, the rocket was simulated at standard temperature and pressure (70 degrees F

and 1 atm) with no launch rail angle. Figure 24 shows the full-scale flight profile of the rocket

under these conditions. The maximum altitude that was predicted was 5,005 feet. Ballast was

still considered for the first flight based off of the baseline simulation because stability was a

concern for the rocket. Another concern was the accuracy of the simulation. A few simulations

before the recorded simulation, the apogee was around 5,600 ft which brought some concern for

the believability of the software. Because of the high apogee, on the simulation before the

baseline, the 2 lb of ballast was used for the baseline and the first flight.

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Figure 24 - Full-Scale Simulation

To make the other flight simulations more like the actual launch day, Figure 25 shows the

launch day flight conditions. These conditions were applied to all three flights that were flown on

the launch day for the full-scale launch.

Figure 25 - Flight Simulation Input Data

0

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4000

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0 20 40 60 80 100 120 140

Alt

itu

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(ft)

Time (s)

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The launch rail sat in the base at an angle of -5 degrees because of the base not being fit

for the University of Evansville’s rail.

For the first flight simulation, a configuration of large quick links (in the recovery section)

and a 2-pound ballast was used for the flight. This configuration was used to give a baseline on

how to modify the rocket for the following flights. For the first flight, a maximum actual apogee

of 4967 feet was reached. The low apogee prompted Project ACE to remove ballast for the

second flight. The low apogee could have been because of the weight from the ballast and the

heavy quick links, or the launch angle. Figure 26 shows the flight profiles for all three different

simulations for the different configurations.

Figure 26 - Simulated Flight Configurations

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0 20 40 60 80 100 120 140

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Flight 2

Flight 3

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The second flight simulation reached an altitude of 5,322 feet, which was within the team’s

goal of a range between 5,200 and 5,400 feet. However, when this configuration was launched,

the rocket came off of the rail at about a 14-degree angle. The angle was because of overlooking

the stability margin of the rocket coming off of the rail after changing the quick links. When the

team removed the ballast from the nosecone, the remaining added weight of the rocket brought

the stability of the rocket below 2 calipers off the launch rail.

With the second flight showing the team that mass was needed in the nose cone, the third

configuration of the smaller quick links and 1.5-pounds of ballast were used in the flight. The

third flight simulation reached an altitude of 5,326 feet. The apogee is within the goal the team

wished to achieve as a first-year team.

The motor thrust curve is given in Figure 27. Appendix D has all component weights for the

different flight configurations.

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Figure 27 - Anticipated Motor Thrust Curve from OpenRocket

One of the NASA requirements was for the rocket to have a minimum rail exit velocity of

52 feet per second. The goal for the team was to have a rail exit velocity of at least 60 feet per

second. The difference in being about 8 feet per second higher than the requirement was to

mitigate the risk of falling below. Using the same simulations as for the previously described,

Table 4 has the predicted rail exit velocities for each flight.

Table 4 - Rail Exit Velocity on Different Flights

Time to Exit Rail

(s)

Velocity at Rail Exit

(feet per second)

Simulation 1 0.44 64.5

Simulation 2 0.43 66.9

Simulation 3 0.43 66.9

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Based on

Table 4, the rail exit velocity is well above the requirement and the team goal. Because the

first flight was the heaviest, the rail exit velocity was lower than the other two flights. Based on

the data, the rail exit velocity for the rocket will be 66.9 feet per second.

One of the other requirements for NASA was the Mach number being less than 1. The

team again, set a goal of being well below the NASA requirement. The team goal was being

below a Mach number of 0.6.

Table 5 - Mach Number on Different Flights

Mach Number

Simulation 1 0.50

Simulation 2 0.53

Simulation 3 0.53

Table 5 shows the predicted Mach Numbers for each of the full-scale flights. Based on the

data in the table, the team goal was met being well below a Mach Number of 0.6. Again, because

the first flight was the heaviest configuration, the Mach Number would be lower. Based on the

flight simulations, the Mach Number of the rocket is 0.53.

Another factor that impacts altitude is the wind speed. Using the same flight conditions, five

simulations were conducted with varying wind speeds from 0 to 20 miles per hour. Table 6

shows the change in the altitude at varying wind speeds.

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Table 6 – Impact of Wind Speed on Altitude

Third Flight Configuration

Wind Speed

(mph)

Predicted Altitude

(ft)

0 5,290

5 5,327

10 5,334

15 5,316

20 5,297

The change in wind speeds plays an important part in the altitude of the rocket. There is a

change in height of about 50 feet due to the variance in the wind speed. Based on the team’s

rocket design, a wind speed of 0 miles per hour would be preferred, while all the wind speeds

allow the rocket to be in the range of the team’s altitude goal.

Validity Assessment

An in-depth analysis comparing subscale flights to OpenRocket simulations can be seen in

the CDR. It was determined that there was a 5% percent error between OpenRocket and actual

flight data. A similarly thorough approach to measuring component weights and dimensions was

used for the full scale simulations. A full list of component weights can be seen in Appendix D.

Figure 28 through Figure 30 graphically shows the OpenRocket and actual flight data for the full

scale flights. A description of the differences and error between the simulations and actual flight

data can be seen in the Flight Analysis section. Table 10 (located in the Flight Analysis section)

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compares the predicted and actual apogees for all three flights. Lastly, in regards to Figure 30

and Flight 3, the main parachute deployed shortly after the drogue parachute. The early ejection

accounts for the large discrepancy between the OpenRocket Simulation Data and the Actual

Data. A further explanation of this can be found in the Flight Results section.

Figure 28 - Flight 1 Actual vs OpenRocket Data

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0 20 40 60 80 100 120 140 160

Alt

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(ft

)

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Actual Altitude (ft)

OpenRocket Altitude (ft)

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Figure 29 - Flight 2 Actual vs OpenRocket Data

Figure 30 - Flight 3 Actual vs OpenRocket Data

0

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Pre-flight, two simulated vehicle factors were validated using empirical data. First, actual

stability was measured in order to assess the validity of the OpenRocket stability values (further

information can be seen in the Actual Stability Margin section). Second, the launch vehicle was

weighed in full to validate final OpenRocket vehicle weight.

Due to the low percent difference in predicted altitude for Flights 1 and 3 (approximately

1%), few changes to the predictive models were made post-flight. Coefficient of drag for the

drogue and main parachutes were empirically determined from actual flight data (further

information can be seen in the Coefficient of Drag section). Acceleration data from the Altus

Telemega was used to empirically calculate the forces acting on the launch vehicle during

parachute deployments (further information can be seen in the Testing section). Launch day

conditions were also used to increase the validity of the flight simulations. See the “Flight

Analysis” section for more detail on the accuracy of the simulation.

Actual Stability Margin

Stability is a metric (measured in calipers) used in rocketry to help determine a rocket’s

ability to maintain its speed and direction. This makes stability vital in designing and testing a

rocket. When considering stability, NASA dictates a minimum stability of 2 cal to ensure that the

rocket would be stable during flight to maintain constant velocity to the target altitude of one

mile. In solving for the stability factor, the following equation was used:

𝑆𝑡𝑎𝑏𝑖𝑙𝑖𝑡𝑦 =(𝐶𝑝−𝐶𝑔)

𝐷 (1)

In this equation, 𝐶𝑝 is the Center of Pressure, 𝐶𝑔 is the Center of Gravity, and D is the

diameter of the body tube of the rocket. The diameter of the body tube is 5.5 inches, and the

45 | P a g e

Center of Pressure is a value determined by simulation from Open Rocket. 𝐶𝑔 is a value that

changes for each flight configuration and has to be determined separately each time the weight in

the rocket is shifted (i.e. a ballast is added to the nosecone). The 𝐶𝑔 for each flight configuration

was determined by hanging the rocket by a rope and balancing. 𝐶𝑔 for each flight configuration

is located at the balance point. This data can be found in Table 7 (The 𝐶𝑝 and 𝐶𝑔 are both

measured from the tip of the nosecone).

Table 7 - Actual Stabilities

𝑪𝒈 (inches) 𝑪𝒑 (inches) Stability (cal)

Flight 1 68. 65 84.3 2.85

Flight 2 71.74 84.31 2.29

Flight 3 69.47 84.31 2.70

The rocket was test launched three times and each time a static stability of above 2 was

calculated, which was above our minimum objective. This shows that the rocket should be stable

in good launch conditions. A sketch of the rocket showing the locations of the 𝐶𝑝 (in red) and

𝐶𝑔’s (in blue).

Figure 31 - Actual Cp and Cg locations

46 | P a g e

Kinetic Energy

It is crucial to ensure that the kinetic energy of the launch vehicle is managed throughout

flight, especially during the final descent. The launch vehicle reaches its maximum kinetic of

173,100 ft-lbf during the ascent, just before motor burnout. To reduce kinetic energy through the

initial descent, the drogue parachute is deployed at apogee and achieves a predicted initial

descent rate of 76.5 ft/s. This gives the heaviest section a kinetic energy of 1249 ft-lbf during the

initial descent.

Upon landing, the kinetic energy of any section of the launch vehicle cannot exceed 75 ft-lbf.

The kinetic energy of each section at landing can be predicted using the mass of each section and

the vehicle’s final descent velocity as predicted by an OpenRocket simulation. These predicted

values are shown in Table 8. The maximum kinetic energy upon landing is 41.0 ft-lbf, which is

experienced by the nose cone and payload.

Table 8 - Predicted kinetic energy of launch vehicle sections

Section Kinetic Energy (ft-lbf)

Nose Cone & Payload 41.0

Coupling Tube 10.88

Booster 33.9

Drift

In order to predict the drift distance of the launch vehicle at landing, five OpenRocket

simulations were conducted for wind speeds of 0, 5, 10, 15, and 20 mph. For each simulation, the

launch angle was set to zero degrees. The resulting drift distances are shown in Table 9. These

47 | P a g e

results verify that the launch vehicle will meet the requirement of limiting drift distance to no

more than 2500 ft even for high wind speeds.

Table 9 - Predicted drift distance for selected wind speeds

Wind Speed (mph) Lateral Distance (ft)

0 9

5 299

10 640

15 1043

20 1492

Full Scale Flight

Launch Day Conditions

The full-scale launch took place at Elizabethtown, Kentucky on Saturday, February 18th.

It was overcast with a chance of rain throughout the day. It was average wind speeds between 4-

8 mph with the cloud layer changing altitude during the day also. The temperature and wind

speed changed throughout the day because of an incoming rain shower. The temperature fell as

the day progressed, however, only the first flight temperature and wind speed was recorded. For

the first launch, it was 59 degrees F at 1 atm pressure with high cloud layer altitude. For the

second launch, the weather conditions changed. It started to rain, but not heavy enough for the

launch to be cancelled. The rain was believed to have an effect on the rocket, but the result of the

effect was uncertain at the time of the launch. The rain could have weighed down the rocket

lowering the altitude, and the humidity could have also caused a change in the actual apogee. For

the last launch, the rain had stopped, but the cloud layer altitude dropped.

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Flight Analysis

Comparison with Prediction

Three flights were completed for the FRR. A summary of these flights can be seen in Table

10. Flight 1 will be used for the OpenRocket prediction analysis. Flight 2 resulted in an unstable

flight with a maximum tilt of 41°. For this reason, Flight 2 was not used for the prediction

analysis and will not be flown at competition. Flight 3 resulted in the main parachute being

deployed prematurely near apogee. For this reason, Flight 3 was not used for the prediction

analysis. Each of the flights will be discussed in further detail in the Flight Results subsection.

Table 10 - Actual Flight vs Predicted Flights Summary

Overall Weight

(lb)

Ballast

(lb)

Predicted Apogee

(ft)

Actual Apogee

(ft)

Percent

Difference

(%)

Flight 1 38.5 2 4,967 4,913 1.09

Flight 2 36.5 0 5,322 4,795 10.42

Flight 3 36.5 1.5 5,326 5,291 0.21

A plot of the actual and predicted altitudes for Flight 1 can be seen in Figure 28 on page 42.

Graphically, it can be deduced that the actual and predicted flight were very similar.

Unfortunately, differing time steps do not allow a direct percent error (Equation (2)) comparison

between the actual and OpenRocket flights. To counteract this issue, a 6 part piecewise

regression line was created based on the actual flight data. This regression line was then

evaluated on the time step of the OpenRocket flight. Error between the best fit line and the actual

49 | P a g e

flight data was calculated, as well as error between the best fit line and the OpenRocket flight

data.

% 𝐸𝑟𝑟𝑜𝑟 =|𝑡ℎ𝑒𝑜𝑟𝑒𝑡𝑖𝑐𝑎𝑙−𝑒𝑥𝑝𝑒𝑟𝑖𝑚𝑒𝑛𝑡𝑎𝑙|

|𝑡ℎ𝑒𝑜𝑟𝑒𝑡𝑖𝑐𝑎𝑙|× 100% (2)

Figure 32 - Actual Altitude vs OpenRocket Altitude

Actual altitude and predicted regression altitude are plotted on Figure 33. Figure 33 also

displays the percent error between these altitudes. The percent error assumed the actual flight

data as the accepted value and the regression data as theoretical. Percent error remained below

11% between 0.55 seconds and 124 seconds. This is the maximum domain that the regression

may be used for when comparing with the OpenRocket data.

0

1000

2000

3000

4000

5000

6000

0 20 40 60 80 100 120 140 160

Alt

itu

de

(ft

)

Time (s)

Actual Altitude (ft)

OpenRocket Altitude (ft)

50 | P a g e

Figure 33 - Actual Data vs Regression

OpenRocket altitude and predicted regression altitude can be seen graphically on Figure 34.

Figure 34 also displays the percent error between these altitudes. The percent error assumed the

OpenRocket data as the accepted value and the regression data as theoretical. Percent error

remained below 11% between 0.55 seconds and 114 seconds.

0%

2%

4%

6%

8%

10%

12%

0

1000

2000

3000

4000

5000

6000

0 20 40 60 80 100 120 140 160

Pe

rce

nt

Erro

r (%

)

Alt

itu

de

(ft

)

Time (s)

Regression Altitude (ft)

Actual Altitude (ft)

Percent Error

51 | P a g e

Figure 34 - OpenRocket Data vs Regression

Although Figure 33 and Error! Reference source not found. show error, it should be

remembered that this is not error between actual and OpenRocket data but rather error between

these and the regression line. It can be concluded from Figure 33 and Error! Reference source

not found. that the largest errors occur at liftoff, main parachute deployment, and low level

turbulence. As this is consistent between both figures, it can be concluded that this large is error

exists due to the regression used to bridge differing time steps. From this, Project ACE has

decided to accept the OpenRocket simulations as a valid prediction method.

Error

The sources of error can be separated into 4 major types. First, there is the inherent error in

the modeling software. Both OpenRocket and Rocksim have documented error within the

program that does not allow for perfectly accurate predictions. To counter this, both programs

0%

2%

4%

6%

8%

10%

12%

0.00

1000.00

2000.00

3000.00

4000.00

5000.00

6000.00

0 20 40 60 80 100 120 140

Pe

rce

nt

Erro

r (%

)

Alt

itu

de

(ft

)

Time (s)

Regression Altitude (ft)

Percent Error

52 | P a g e

are used, in order for each to validate the other. Error within the programs was discussed

extensively in the PDR.

Secondly, there is systematic error in inputs to the modeling software. For example, the

accuracy of lengths is limited to the accuracy of the ruler used to measure them. Alternatively,

some parameters could not be measured and were thus based on research. For instance, the

surface roughness of carbon fiber was not measured, but was instead based on research. Third,

there is random error. Similar to the fluctuation of a needle on a gage, there will be a certain

variance in the apogee of the rocket from one flight to the next.

Lastly, there is error in the best fit curve created to compare OpenRocket data to actual flight

data. This is mentioned and described in the previous section.

Coefficient of Drag

The coefficient of drag is simulated by OpenRocket from liftoff until drogue deployment. A

plot of this can be seen in Figure 35. At drogue deployment, the coefficient of drag is assumed to

be equal to the manufacturer specified coefficient of drag of the drogue parachute. At the time of

the main parachute deployment, the coefficient of drag is assumed to be equal to the

manufacturer specified coefficient of drag of the main parachute.

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Figure 35 – Predicted Coefficient of Drag During Flight

Coefficient of drag was calculated based on experimental values. The Atlus Telemega

altimeter located in the nosecone of the launch vehicle records acceleration and velocity data.

Post motor burnout, only drag and weight act on the launch vehicle. Using summation of forces,

drag can be calculated using the following equation (where acceleration and gravity both act in

the positive direction):

𝐷 = 𝑚(𝑎 − 𝑔) (3)

Following the calculation of drag force, coefficient of drag can be calculated using the

following equation:

𝐶𝐷 =𝐷

𝜌𝑣2

2𝐴𝑐

(4)

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0

1000

2000

3000

4000

5000

6000

0 5 10 15 20

Co

eff

icie

nt

of

Dra

g

Alt

itu

de

(ft

)

Time (s)

Altitude (ft)

Drag coefficient

54 | P a g e

These equations are valid post motor burnout and pre-drogue deployment. An average of

acceleration and velocity were taken over this range. The experimental coefficient of drag was

calculated to be 0.397. Compared to the average OpenRocket coefficient of drag over the same

range (0.449), this is a 12% difference.

Flight Results

The team launched three times with success on each of the launches. Table 10 shows the

apogee results from each of the three flights. For the first flight, the team utilized a launch rail

provided by the University of Louisville. The launch vehicle was equipped with 2 lb of ballast,

and was mostly successful; the vehicle came straight off the launch rod, recovery events

occurred at the correct times, and no damage was observed. However, the recorded apogee of

4913 ft was well under the team’s minimum goal of 5200 ft. The first flight at the launch site

showed the first simulation and the baseline simulation were correct on OpenRocket. The flight

with the 2 lb of ballast was ran because of the worry that the OpenRocket simulation was

incorrect. The reasoning behind the worry of the simulation not being correct was because a few

simulations before the final, the apogee was around 5,600 ft. To mitigate any worry with the

simulations being incorrect, the 1st configuration was ran for a starting point of the ballast

optimization and to double check the OpenRocket simulations.

In order to increase the apogee of the second flight, the 2 lb of ballast were removed. The

second launch was less successful with only the recovery being a success. The problems that

occurred in this flight were the altitude being too low, the drift distance was too far, and the

stability too low. The second flight reach an altitude of 4,795 feet, which was lower than the first

flight. The reason for the low altitude was an unexpectedly low stability. When the ballast in the

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nosecone was removed from the rocket, the center of gravity was lowered toward the aft of the

rocket lowering the stability. The rocket came off of the launch rail at an angle of 14 degrees

from vertical and ended around 40 degrees from vertical a few hundred feet above the ground.

The angles were acquired from the altimeter located in the nosecone. The angles led to a lower

apogee because of the trajectory of the rocket traveled. Also, due to the angle that the rocket

launched, it landed further way from the launch site and in a tree. No parachutes were torn or any

part of the rocket harmed when it landed in the tree or when it was removed.

The third configuration flown was 1.5 lb of ballast with the smaller diameter quick links

mentioned in the “Changes Made to Vehicle Criteria” section. The last configuration shifted the

center of mass further toward the bow, increasing the stability of the launch vehicle to fix the

issues observed during the second flight. The third flight was the most successful with respect to

the apogee achieved, however unexpected performance of the recovery system resulted in a large

drift distance. The apogee of 5291 ft satisfied the team objective to reach within 200 feet of one

mile. However, the main parachute deployed early, just after deployment of the drogue

parachute. This resulted in a velocity of 15 ft/s for the entirety of the descent. The wind then

carried the launch vehicle to just over one mile from the launch site. While the drift distance was

greater than the acceptable maximum, the vehicle was able to be recovered without damage.

The early deployment of the main parachute was likely due to over-packing of the ejection

charges for the third flight; the scale available on-site was not as precise as the one used to

measure the ejection charges for the first two flights which was prepared in advance and as a

result a larger amount of black powder was used. The larger charge likely caused the bow body

tube to separate at a high velocity, pulling the coupling tube out of the aft body tube prematurely.

56 | P a g e

Despite the excessive drift distance of the third flight, the team has selected this

configuration to be flown at competition. The ejection charges will be measured precisely prior

to the competition to ensure that the recovery system performs as expected. The recorded apogee

for this configuration should fulfill the team’s goal.

Payload Criteria

After initial drop testing proved that neither welding nor epoxying the springs to the

bullheads would suffice, the design was changed to a bolt and washer mounting assembly seen in

Figure 36.

Figure 36 - Final Design Assembly (new bolt and washer mounting)

The design change used 30 washers and bolts threaded into both bulkheads to secure the

bottom and top layers of each base spring. Each spring had 3 washers on either side allowing

one to tighten or loosen one of the bolts to assure the spring was at a constant 90-degree angle to

avoid buckling. Due to the addition of 30 bolts, the team repurposed the old bulkheads by

adding 15 more threaded holes to each. The drawing of each bulkhead can be seen in Appendix

A. The exploded view of the entire payload assembly can be seen in Figure 37 focusing on the

base spring assembly.

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Figure 37 - Exploded view of payload assembly (annotation following)

(1) Represents the U-bolt that is attached to the main parachute which screws into the

recovery bulkhead (3). This is epoxied directly to the ID of the rocket’s main body tube. The

recovery side bulkhead for the payload (5) is attached to the recovery bulkhead (4) via two bolts

shown in the figure as (2). (4) is a clear spacer to separate the two recovery side bulkheads. (6)

shows the 30 washers used to hold the 5 base springs, labeled (8), in place by inserting them

above the last two coils in each spring. (7) is the 30 bolts used to tighten the washers, (6), and

base springs, (8), into place. (9) shows the bulkhead epoxied in Cylinder 2, (10). Finally, (11)

shows Cylinder 1, a 3D printed canister mounted within Cylinder 2, (10) via the 12 CR1-400

wire rope isolators labeled (12). A second view of the exploded assembly showing how the

washers and bolts attach the base springs to each bulkhead can be seen in Figure 38.

Figure 38 - Exploded view base spring attachment

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Another design change that was implemented due to a failure during initial testing was the

addition of thin aluminum squares epoxied to Cylinder ,1 as well as pins used in all mounting

points of the wire rope isolators. During initial tests, some of the epoxy failed due to shear

stress, causing the wire rope isolators to break free from the ID of Cylinder 2. The solution

applied was to drill holes into cylinder 2 and epoxy pins inserted in the thru hole of each end of

the spring. Both cylinders had holes were drilled to add more strength and reduce total shear

stress felt by the epoxy. This solved the epoxy’s adhesive failure, but Cylinder 1 experience 2

cases of cohesive failure where the 3D printed plastic was ripped apart due to a weakness in

tension. To combat this, thin .1-inch-thick aluminum squares 1x1 inch were epoxied to Cylinder

1 to spread the force over several layers of material. The wire rope isolator with epoxied pins and

aluminum plate can be seen in Figure 39.

Figure 39 - CR1-400 wire rope isolator pin and plate assembly

(1) shows the 0.1-inch thick aluminum plate used to disperse the force along several layers of

the 3D printed plastic of Cylinder 1. (2) shows the CR1-400 wire rope isolator and (3) shows

epoxied pins in the thru holes of the isolator that would go on to be inserted into Cylinders 1 and

2. The pins were cut from a standard carpentry nail. The pins used in attaching the wire rope

59 | P a g e

isolators not only add strength by transmitting some of the shear force into the pin, but make

assembly easier by fitting the pins into holes in both Cylinders 1 and 2.

Safety

The University of Evansville’s first and foremost priority throughout the duration of this

project has been and will continue to be a focus on safety. This consideration and team-wide

emphasis on safety has been paramount in this project, as it has allowed the UE’s SLI team to

stay on schedule and create a safe and successful launch vehicle. Throughout the duration of this

project, in order to create the safest possible working and testing atmosphere, risks were

identified and mitigations were developed before material handling, fabrication operations, or

testing was completed. In addition to this, all team members have been, and will continue to be,

educated on the risks associated with all areas of the project. This is significant because,

education allows team members to fully understand the risks associated with the operations/items

that team members are coming in contact with, and details the proper procedure to take in order

mitigate these risks.

In the following tables, various hazard and failure mode analyses of the launch vehicle will

be considered in order to present possible risks associated with the project, and detail mitigation

tactics and verification plans that will be used to alleviate these risks. In order to generate these

continually updated tables, the team first began by brainstorming the possible risks associated

with each individual section of the rocket from fabrication, to handling of materials, and launch

operations. As the project progressed from the design phase to the fabrication phase, and

ultimately to the testing phase, the team was able to further identify other unforeseen risks as

60 | P a g e

well as develop and conduct verification tests in order to mitigate various risks. In the hazard and

failure modes analysis tables, the impact and likelihood of each risk was assessed and quantified

using the definitions provided in Table 11.

Table 11 - Definitions for Hazard and Failure Mode Analyses

Severity Definition

1-Catastropic Extreme reduction in safety; potential complete loss

2-Critical Substantial reduction to overall safety or functionality

3-Marginal Minor reduction to overall safety or functionality

4-Negligible Little to no reduction in overall safety of team members or

component functionality

Likelihood Definition

A-Frequent Occurrence of the event is expected

B-Probable Occurrence of event is likely, but not guaranteed

C-Occasional Chance of occurrence is possible, but not significant

D-Remote Minor change of occurrence

E- Improbable Occurrence of event is extremely unlikely

Following this categorization, mitigations and verification plans were proposed in order to

decrease both the significance of the risk as well as the change of occurrence. Lastly, the risk

was then reevaluated in order quantify the impact of the mitigations methods.

Personnel Hazard Analysis

A personnel hazard analysis was conducted to identify hazards, effects, likelihood of

occurrence, and impact of individual factors associated with project. Safety practices and

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protocols were created to make team members aware of potential hazards, and reduce the chance

of risk or injury during the course of the project. The personnel hazard analysis is summarized in

Table 12 through Table 16.

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Table 12 - Personnel Hazard Analysis - Epoxy

Risk/Hazard Root Cause/Effect Severity/

Likelihood Mitigation and Control

Post-Control

Severity/Likelihood

Epoxy Fumes

Open containers of epoxy during

fabrication operations leading to

inhalation of toxic fumes,

accidental ingestion, or contact

with skin leading to potential for

irritation or rash

4A Work in well ventilated spaces 4C

Epoxy

Contacting Skin

Mishandling of epoxy during

application leading to skin

irritation

4A

Individuals handing epoxy must be wearing

Proper PPE, such as gloves, pants, and close-

toed shoes when handling epoxy to prevent

contact with the skin. In the event that epoxy

does come in contact with the skin, wash it

off at the sink

4C

Spill/Leak of

Epoxy

Mishandling of epoxy resulting in

epoxy hardening on the working

area, potentially ruining lab

equipment or various parts of the

launch vehicle

4C

Handle the epoxy carefully during mixing or

transport. In the event that any epoxy does

spill, wipe up the excess with a cloth and

dispose of it properly, and clean the dirtied

area.

4D

Epoxy Burning

Through

Container

Mishandling of epoxy leading to

potential damage to user, lab, or

equipment

2E

Never leave epoxy unattended. Monitor the

heat of the epoxy as you mix it. If epoxy does

get excessively hot, remove sample from lab

and let it cool before disposing of it properly

2E

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Table 13 - Personnel Hazard Analysis - Launch Operations/Post-Launch Inspection

Risk/Hazard Root Cause/Effect Severity/

Likelihood Mitigation and Control

Post-Control

Severity/Likelihood

Debris In Team

Member's Eye

Particles flying through the air

during fabrication operations

leading to potential scrape or cut

to user's eyes

2C

Wear proper PPE, such as safety glasses

during launch and fabrication. In the event

that debris does enter the eye, the eyewash

station will be used to cleanse the eye of the

debris.

3D

Sharp Edges on

Fins and

Nosecone

Improper sanding or fabrication of

fins and nose cone resulting in

potential splinters or cuts to team

members

4D

Team member will be required to wear proper

PPE, such as gloves, close-toed shoes, and

pants during testing operations and inspection

procedures to prevent direct contact with

fragments of the rocket

4E

Cracks or

Chipping in

Body Tube

Improper fabrication operations or

faulty components putting team

members at risk for splinters or

cuts when coming in contact with

rocket

3C

Team member will be required to wear proper

PPE, such as gloves, during testing setup and

inspection procedures to prevent direct

contact with fragments of the rocket

4E

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Materials

Experience

Explosive

Breaking When

Opening for

Inspection

Failure of component durability or

subsystem resulting in a range of

possible injuries to team members

from minor to severe depending on

the intensity of the explosion

1E

Team members will wait in designated safe

launch zone until rocket is deemed safe for

retrieval by RSO. Safety officer will retrieve

rocket, wearing proper PPE and keep face

directly out of line of launch vehicle.

2E

Direct Contact

With Hot

Material

Oversight or ignorance when

approaching hot materials for

handing, yielding to varying

degree of burn to team members

2D Proper PPE, such as gloves or aprons, must be

worn at all time when handling hot objects 4D

Materials

Catching Fire

Improper storage of flammable

components or inappropriate

fabrication operations/tool usage

leading to potential severe injury

or burns to team members,

equipment, or work space

1D

All flammable objects will be kept in proper

locations away from sparks and open flames.

In the event of a small fire, a fire extinguisher

will be used to put out the fire. In the event of

a large fire, the team will evacuate the

building and the fire department will be called

2E

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Black Powder

Fumes

High exposure to black powder

when handling and preparing

samples of this toxic gas can result

in coughing, dizziness, and

fainting

2D

Black powder is stored in portable fireproof

case to keep away from fire and high

temperatures. When handling substance

recovery subsection lead will measure

samples in well ventilated areas

3D

Rocket

Propellant

Comes In

Contact With

Skin

Improper transportation and

configuration of motor subsystem

leading to irritation and burns

2C

Per the motor preparation checklist, the motor

will be transported from an offsite location to

the launch location in a protective, waterproof

casing. Upon installation, the propulsion team

lead will prepare the motor according to

manufacturer specification while wearing

proper clothing, shoes and PPE

2D

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Table 14 - Personnel Hazard Analysis - Testing

Risk/Hazard Root Cause/Effect Severity/

Likelihood Mitigation and Control

Post-Control

Severity/Likelihood

Electrical Shock

Unsafe working conditions or lack

of care when handling electronics

leadings to electrocution resulting

in burns, significant injury, or

death

3D

Per malfunctioning electronics

troubleshooting checklist, electronics

subsection lead will inspect faulty instrument

for improper connection. Care has been take

to ensure nothing with exposed or fraying

wiring is being used in fabrication.

Electronics will be stored in dry, secured area

3E

Inexperienced

Test Personnel

Improper handling of shop tools or

machining operations leading to

personal injury or destruction of

equipment

3C

Only authorized individuals have run tests.

Multiple team members are present during

testing to report and issue if one should occur

3D

Fractured

Particles During

Testing

Failure of various components

leading to potential splinters or

cuts to team members

3B

Team member have been required to wear

proper PPE during testing setup and

inspection to prevent direct contact with

fragments of the rocket

4E

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Table 15 - Personnel Hazard Analysis - Fabrication

Risk/Hazard Root Cause/Effect Severity/

Likelihood Mitigation and Control

Post-Control

Severity/Likelihood

Allergic

Reaction to

Building

Material

Handling of materials team

member is allergic to resulting in

an allergic reaction in the form of

skin irritation, rash, or swelling

2E

Proper clothing, shoes, and PPE must be worn

at all time when handling materials. Allergies

of all team members are kept on file, and

members allergic to a specific material will

not work with that material while it is

fabricated.

2E

Improper Heavy

Machinery

Usage

Improper handling of shop tools or

machining operations leading to

personal injury or destruction of

equipment

2C

All team members have been trained on how

to properly use shop equipment and have

passed written and practical tests regarding

proper handling and maintenance of shop

equipment

2D

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Improper

Handheld Tool

Usage

Bruises, cuts or scrapes from

mishandling of basic handheld

shop tools such as hammer or saw

3C

All team members have been trained on how

to properly use handheld tools. During

fabrication operations, team members have a

spotter to ensure proper safety procedures are

followed and to monitor surroundings during

fabrication operation

3D

Improper Tool

Storage

Tool storage in improper location

following fabrication operations

leading, or usage by unauthorized

individuals leading to damage to

equipment or environment.

3C

Tools are stored in proper locations to keep

team members and work area clean and safe,

and prolong life of tool. Periodic checks will

be conducted by safety officer to ensure all

materials are returned following construction

and placed in their proper locations

3D

Improper Use

of Craft/Exacto

Knife

Cuts leading to injury as a result of

unsafe precision cutting operations

on fins or other pieces of the

rocket body

2D

During fabrication operations, team members

have at least one spotter to ensure proper

cutting procedures are being followed by

cutting away from body.

2E

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Improper Work

Attire

Lack of education on proper

clothing or inspection of shop

workers leading to potential

damage to body or attire

4D

Proper clothing, shoes, and hair styles will be

required in the lab to ensure safety for all

team members. Safety officer will conduct

periodic checks of fabrication work attire and

PPE.

4E

Tripping

Hazards

Cords or other materials lying on

the floor could cause team

members to trip, thus resulting in

cuts, scrapes, bruises, or broken

bones

4B

Cords will be plugged in closest to the area in

which their machine is being used. The work

area will be kept tidy in order to prevent

debris from accumulating on the floor

4D

Overreaching

Lack of awareness to surroundings

leading to potential falls, cuts, or

scrapes

4B

Ensure all needed materials are close by

before beginning fabrication to avoid

overextension. Keep proper footing/balance.

4C

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Table 16 - Personnel Hazard Analysis - Education Engagement Outreach Events

Risk/Hazard Root Cause/Effect Severity/

Likelihood Mitigation and Control

Post-Control

Severity/Likelihood

Car Accident

Not following driving rules and

regulations resulting in a range of

potential injuries to team members

in the car, or others from minor to

severe, and potential property

damage

1E

Seat belts are worn at all times by all

members inside the vehicle. All individuals in

the vehicle will also sign a waiver releasing

the team of liability in the event of an

accident. The driver must follow all federal

driving laws including have a valid license

and insurance

2E

Child Using

Tools

Inappropriately

Lack of oversight of individuals

managing event or disobedient

participants could lead to child

experiencing a range of injuries

depending of the tool being used at

the event

2D

Age appropriate tools will be given during

educational outreach events. Strict

supervision will be used to monitor all

activates to ensure all children are safe and

know what they are supposed to do.

3D

Child Not

Following

Instructions

Lack of oversight of individuals

managing event or disobedient

participants could lead to child

could experience a range of

injuries depending operation/event

3D

All children will be closely monitored in

order to ensure they are doing what they are

supposed to. If they continue to be defiant,

they will be removed from the activity.

4D

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Failure Modes and Effects Analysis

In order to analyze the functionality and safety of the rocket and all of its components, a

failure modes and effects analysis was created. In this analysis, presented in Table 17 through

Table 23, verification plans, referencing various pre-launch checklists or data obtained from tests

conducted on individual components, are stated in order to verify mitigation tactics to reduce

risks. Then, post-control severity and likelihood was then reevaluated to see the impact that the

mitigation tactics and verification checks had on the risk.

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Table 17 - Failure Modes and Effects Analysis - Design/Fabrication

Risk/Hazard Root Cause/Effect

Severity/

Likelihood

Mitigation and Control Verification Plan

Post-Control

Severity/Likelihood

Cracking or

Chipping of

Fabricated Parts

Faulty components or

inappropriate fabrication

operations lead to failure of

rocket upon launch or testing

operation. Potential to splinter

and cause significant damage to

other sections of the rocket or

lead to failure of subsequent

components

2D

Care has been taken to ensure all

parts have been fabricated

according to specification. Parts

will be stored in appropriate

containers and holders within

locked room to prevent

accidental damage.

In accordance with the final

assembly checklist, each

subsection lead will inspect their

section of the rocket for any

compromises in structural

integrity as a result of fabrication

operations

3D

Lack of Precision

When Fabricating

Parts

Fabrications not completed

according to specifications

leading to potential inability to

assemble components of rocket

properly and have secure

attachment, resulting in failure

of rocket upon launch or testing.

2D

Only trained individuals are

allowed to operate any

machinery during the fabrication

and construction process. Other

team members will verify work

to ensure it meets standards set

for in the design

In accordance with the final

assembly checklist, each

subsection lead will inspect their

section of the rocket to ensure all

parts are fabricated to specified

dimensions

3D

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Gaps Between

Connecting Pieces

Fabrications not completed

according to specifications

resulting in inability to assemble

components of rocket properly

and have secure attachment

potentially leading to failure of

rocket upon launch or testing

operation.

2C

All components of the rocket

have been measure after

fabrication in order to ensure

they meet the dimensions

specified in the design. The pre-

launch safety checklist will be

used to ensure team members

visually inspect connections of

components prior to launch

In accordance with the final

assembly checklist, each

subsection lead will inspect their

section of the rocket to ensure all

parts are fabricated to specified

dimensions. In the event that

there are gaps between adjoining

sections, the troubleshooting

checklist will be followed to

remedy the issue

2D

Insufficient Epoxy

Lack of attention to security of

connection causing inability of

components to hold together

leading to separation and

potential failure

3D

Epoxy has been mixed in

accordance with instructions in

order to ensure a good adhesive

mixture. Components will be

tested prior to launch to ensure a

secure, water-tight seal

In accordance with the secure

attachment inspection within the

final assembly checklist, each

subsection lead will inspect their

section of the rocket to ensure

attachment between adjoining

subsections

3E

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Wrong Equipment

Usage for

Fabrication

Operation

Lack of knowledge by team

members during fabrication

operations leading to potential

damage to component(s) of

rocket and individual harm. Also

potential to generate flawed

component that is not suitable

for usage

4D

Only trained individuals are

allowed to operate any

machinery during the fabrication

and construction process. If a

component is corrupted,

fabrication will be done to

salvage as much of the material

as possible without

compromising safety of the

launch vehicle and operations

All team members will pass a

practical and written test on

proper usage of shop equipment

before they are allowed to use

equipment.

4E

Materials Catch Fire

Improper storage or handling of

materials causing potential

damage to component(s) of

rocket and individual harm

resulting in minor to major loss

of equipment, workspace, or

components, or compromise of

structural integrity of launch

vehicle

1D

Team members will operate in a

safe manner to prevent the

start/spread of fire. In the event

of a small fire, it will be

extinguished using the fire

extinguisher in the energy

systems lab. For larger fires, 911

will be called and the team will

retreat to a safe distance.

Team members will be trained

on how to properly use fire

extinguisher in the event of a

small fire. Safety officer will

periodically test fire extinguisher

to ensure it is fully functional.

2E

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Improper Storage of

Materials/Equipment

Lack of knowledge as to where

to put supplies when fabrication

operation is completed leading

to potential damage to

materials/equipment resulting in

compromise of the structural

integrity of various components,

or inability to use equipment for

further fabrication operations

4A

The energy systems lab is

cleaned after each work period.

Checklists have been created to

ensure that all materials and

equipment being used are

returned to their proper locations

before everyone can leave.

The parts checklists will be used

to sign tools and materials in and

out. Additionally, the safety

officer will periodically check

the supply cabinets to ensure all

tools are returned and in their

proper locations following

fabrication operations

4B

Degradation of

Epoxy

Oversight of connection security

during inspection process or

improper storage leading to lack

of adhesion between parts

resulting in separation and

potential failure

3E

Epoxy is stored in the in the lab

at room temperature according

to specification listed by the

manufacturer.

Parts checklists will be used to

check out tubes of epoxy so that

the safety officer has all supplies

accounted for. Furthermore, each

subsection of the rocket will be

inspected via the final assembly

checklist to ensure proper

connection and adhesion

between adjoining sections.

3E

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Table 18 - Failure Modes and Effects Analysis - Payload

Risk/Hazard Root Cause/Effect

Severity/

Likelihood

Proposed Mitigation Verification Plan Post-Control

Severity/Likelihood

Premature Ignition

Charge

Failure to properly prepare

ignition system according to

checklist leading to a range of

failure modes from minor

damage to the payload system or

components to catastrophic

failure due to premature

separation and parachute

deployment

2E

Black powder is stored in a dry

area at room temperature in

accordance with manufacturer

specifications. Testing has been

done to ensure premature

ignition does not lead to

recovery failure

See ejection testing summary

and results in project plan

section of FRR

3E

Failure of Motor

Faulty motor or improper

storage leading to inability for

rocket to ascend off launch pad.

Potential damage to payload or

other components upon misfire

3D

Testing and research has been

completed in order to ensure the

proper motors for each size of

launch vehicle created is being

used. Rocket motors will be kept

in a dry area at room

temperature in accordance with

manufacturer specifications.

Motor has been tested via full-

scale launch operations. For

further detail see full scale

testing section of FRR

3E

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Failure of Black

Powered Charge

Improper storage of black

powder sample or compromised

sample resulting in an inability

for launch vehicle to separate

leading to potential catastrophic

damage to payload and rocket

failure of recovery system

1D

Black powder is stored in a dry

area at room temperature in

accordance with manufacturer

specifications. Testing has been

done to ensure proper amounts

of black powder is used for

ignition.

See ejection testing summary

and results in project plan

section of FRR

1E

Deployment of

Black Powder

Change Resulting in

Damage to Payload

Holding Container

Inability to input proper amount

of black powder into launch

vehicle as determined by testing

resulting in over pressurized

capsule causing minor to

catastrophic damage to payload,

holding container, or spring-

damper system

1E

Black powder is stored in a dry

area at room temperature in

accordance with manufacturer

specifications. Testing is done to

ensure proper amounts of black

powder is used for ignition.

Ejection test was completed in

order to determine the proper

amount of black powder to be

used to pressurize the launch

vehicle and deploy the

parachutes. Impact tests were

also completed on the payload

container to determine how

much force is felt by the fragile

material while within the

dampening system

2E

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Bending/Breaking of

Spring-Damper

System

Failure to properly account for

forces experienced by launch

vehicle during flight or improper

assembly resulting in

compromise in the structural

integrity of the spring-damper

system and damage to the fragile

payload

3A

Testing has been conducted in

order to minimize forces on

payload and ensure fragile

payloads of all types will be kept

safe and secure during launch

and recovery operations

Tensile and impact tests were

completed on the payload

container in order to measure it's

tensile and ability to dampen

direct impact. For further detail,

see MTS testing summary and

results in project plan section of

FRR

3D

Crack in Payload

Holding Container

Failure to properly account for

forces experienced by launch

vehicle during flight or improper

inspection prior to launch

resulting in compromise in the

structural integrity of the spring-

damper system and damage to

the fragile payload

3D

An inspection has been

completed in accordance with

the pre-launch safety checklist in

order to ensure the structural

integrity of the payload systems

is not in any way compromised

prior to launch operations

In accordance with the final

assembly checklist, the payload

container will be inspected for

cracking or any other structural

imperfections that could have

been acquired during fabrication

or transport prior to launch

3E

Inability to Keep

Payload Static

within Holding

Container

Dampening material is unable to

absorb impact and restrict

movement leading to potential

minor to catastrophic damage to

fragile payload and damage to

other components of the launch

vehicle. Potential failure to meet

mission objective

2D

Testing was completed in which

the movement of the payload

within its holding container is

measured in order to ensure it

does not experience collision

with surrounding walls or

anything else that could cause

fracture or damage.

See MTS testing summary and

results in project plan section of

FRR

3D

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Payload Damage

Upon Impact With

Ground Upon

Decent

Inability to slow speed of launch

vehicle during decent leading to

damage to fragile payload

resulting in repair, replacement,

or failure to meet mission

objective

2E

Tests were conducted in order to

validate that the materials used

for the payload security

container can withstand the

forces experienced by the fragile

material without damage to its

structural integrity

See MTS testing section of FRR

for results on impact testing of

payload container

2E

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Table 19 - Failure Modes and Effects Analysis - Payload Integration

Risk/Hazard Root Cause/Effect

Severity/

Likelihood

Proposed Mitigation Verification Plan Post-Control

Severity/Likelihood

Lack of Space in

Body Tube for

Payload Container

Fabrication operations not

completed according to

specifications leading to inability

to properly assembly launch

vehicle resulting in removal of

payload container and failure to

meet mission objective

3C

Throughout the fabrication

process all complete parts have

been measured and verified by

the team lead in order to ensure

they meet the proper dimensions

laid out in the design. This will

allow for proper fit and connect

within the launch vehicle

In accordance with the final

assembly checklist, each

subsection lead will inspect their

section of the rocket to ensure all

parts are fabricated to specified

dimensions. In the event that

there are gaps between adjoining

sections, the troubleshooting

checklist will be followed to

remedy the issue

3E

Payload Container

Not Properly

Mounted in Body

Tube

Failure to properly inspect

payload subsection system prior

to launch causing potential

damage to the payload or its

housing container. Could

compromise the structural

integrity of various components

or lead to failure of other

operations

2D

Prior to launch the pre-launch

checklist will be used to verify

that payload is mounted

correctly in place and all

connections are secure to ensure

safe launch operations

In accordance with launch

procedures checklist, payload

will be reviewed for flight

readiness and proper mounting

prior to launch operations

3D

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Weak Attachment

Between Payload

Container and

Recovery System

Failure to properly inspect

connection between adjoining

subsections leading to possible

cracking or separation between

the payload and recovery

systems and inability to return

fragile material

2D

Prior to launch the pre-launch

checklist will be used to verify

that payload is mounted

correctly in place and all

connections are secure to ensure

safe launch operations

In accordance with the secure

attachment inspection within the

final assembly checklist, the

connection of adjoining

subsections will be checked by

the safety officer to ensure

proper connection. In the event

that there are gaps between

adjoining sections, the

troubleshooting checklist will be

followed to remedy the issue

3D

Inability to Fit

Given Payload Into

Container

Failure to fabricate payload

container according to

specifications given by NASA

resulting in potential inability to

meet mission objective of safely

launching and returning a fragile

material on our launch vehicle

3E

All fully fabricated parts have

been measure and compared to

design requirements in order to

ensure they meet the proper

dimensions, thus ensuring that

the fragile payload fits within the

envelope of the container

Payload container has been

design in accordance with the

envelope of fragile material

provided by NASA

3E

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Inability to Fill

Payload Container

with Material that

Dampens Force Felt

by Payload

Improper material used to

dampen forces or unreliable

impact testing data causing

damage to the payload that could

result in failure to meet mission

objective

2C

Testing has been done in order

to measure the force felt by the

payload during launch and

landing operations

See MTS testing section of FRR

for results on filler material's

ability to dampen impact

3D

Inability to Fill

Payload Container

with Material that

Restricts Payload

Movement During

Flight

Improper material used to

restrict movement or unreliable

impact testing data resulting in

damage to the payload that could

result in failure to meet mission

objective

3C

Testing has been done in order

to measure the movement of the

payload within the container in

order to ensure it will not be

damaged as a result of striking

interior walls

See MTS testing section of FRR

for results on filler material's

ability to restrict movement

4D

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Table 20 - Failure Modes and Effects Analysis - Recovery System

Risk/Hazard Root Cause/Effect

Severity/

Likelihood

Proposed Mitigation Verification Plan Post-Control

Severity/Likelihood

Parachute is Not

Packed Properly

Improper packing method

leading to failure in parachute to

deploy properly resulting in

launch vehicle experiencing

more force than planned upon

when landing

1E

Testing have been done in order

to verify the packing method and

give the student practice with

packing the parachute into the

body tube. On launch day, the

pre-launch checklist will be

followed to ensure proper

packing of the parachute in

accordance with standard

practices.

Various parachute packing

methods have been researched

and tested in order to determine

an optimal method. Tests have

been conducted with various

packing styles. For further detail,

see parachute deployment force

testing subsection of FRR

3E

Tear in Parachute

Failure to properly inspect

parachute prior to launch or

faulty parachute resulting in

launch vehicle descending at a

faster rate than planned in an

uncontrolled manner causing

potential damage to components

or total loss

2D

Various tests have been

completed in order to verify the

strength of the parachute. The

parachute will be inspected prior

to launch using the pre-launch

checklist in order to verify it

does not have any tears, pulls,

rips, or other imperfections that

could result in failure

In accordance with the recovery

preparation checklist, all

parachutes will be inspected

prior to launch for tears, snags,

or any other imperfection that

could result in recovery failure

3E

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Tear in Shock Cord

Failure to properly inspect shock

cord prior to launch or faulty

component resulting in launch

vehicle descending at a faster

rate than planned in an

uncontrolled manner potentially

damaging components or

resulting in a total loss

2E

Tests have been run in order to

verify the strength of the shock

cords. The cords will be

inspected prior to launch using

the pre-launch checklist in order

to verify it does not have any

tears, pulls, rips, or other

imperfections that could result in

failure

In accordance with the recovery

preparation checklist, all shock

cords will be inspected prior to

launch for tears, snags, or any

other imperfection that could

result in recovery failure

3E

Shock Cord Cannot

Withstand Force of

Parachute

Deployment

Incorrect shock cord for force

experienced during deployment

causing separation of rocket into

multiple pieces, some of which

will not be attached to the

parachute, causing damage and

potential harm to by standards

1D

Testing has been done in order

to verify the strength of the

connection between the shock

cords and the main rocket body

tube. The connection between

these two will also be inspected

prior to launch with the pre-

launch checklist in order to

ensure a secure attachment

See parachute deployment force

testing subsection of FRR for

details regarding shock cord

strength and durability for

various black powder charges

2D

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Drogue Parachute

Deployment Failure

Failure to properly pack

parachute or use correct amount

of black powder for

pressurization leading to

uncontrollable decent until the

opening of the main parachute

resulting in the launch vehicle

landing with a greater impact,

thus causing damage to

components or endangering

spectators

1E

Tests have been conducted in

order to verify the precise

amount of black powder that will

need to be used to pressurize the

parachute and allow for proper

deployment.

See parachute deployment

testing and full-scale testing

subsections of FRR for details

regarding proper quantity of

black powder for launch vehicle

pressurization and drogue

parachute deployment

3E

Main Parachute

Deployment Failure

Failure to properly pack

parachute or use correct amount

of black powder for

pressurization leading to decent

at a quicker rate than expected

and potential drift of the launch

vehicle off course resulting in

damage to the components or

endangerment of spectators.

1E

Various tests have been

completed in order to verify the

precise amount of black powder

that will need to be used to

pressurize the parachute and

allow for proper deployment.

See parachute deployment

testing and full-scale testing

subsections of FRR for details

regarding proper quantity of

black powder for launch vehicle

pressurization and main

parachute deployment

3E

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Wind Blows Rocket

Off Course

Launch of rocket in excessively

windy conditions resulting in an

inability for parachutes to open

at proper altitudes or launch at

excessive wind speeds leading to

potential for rocket to become

lost or components becoming

damaged if landing occurs

outside of the space given.

3D

The rocket will only be launch in

proper conditions, therefore

minimizing the chance for wind

gusts to blow the rocket off

course. In the event that this

does occur, the launch vehicle

will be retrieved using the GPS

tracking system in the altimeter

UE SLI team in conjunction with

NASA and local rocket clubs

will monitor wind speeds and

make launch related decisions

accordingly

4E

Parachute Deploys

at Incorrect Time

Incorrect packing of parachute,

or faulty electronics leading to

potential for uncontrollable

decent, damage to components

or compromises to structural

integrity of the launch vehicle

2D

Testing has been done in order

to verify the precise amount of

black powder that will need to

be used to pressurize the

parachute and allow for proper

deployment at the correct time.

The recovery system will also be

tested prior to launch in

accordance with the pre-launch

checklist.

See altimeter testing and full-

scale testing subsections of FRR

for details regarding proper

quantity of black powder for

launch vehicle pressurization

and parachute deployment at

proper times

3D

Interference from

the Scoring

Altimeter Causes

System Failure

Operating on the same, or close

to the same frequency resulting

in a failure in the recovery

system.

1D

Testing the recovery system with

the scoring altimeter on and

nearby to determine if there is

any interference with the system.

Verify that all electronics are

working according to the Pre-

Launch checklist.

1E

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Table 21 - Failure Modes and Effects Analysis - Testing

Risk/Hazard Root Cause/Effect

Severity/

Likelihood

Proposed Mitigation Verification Plan Post-Control

Severity/Likelihood

Inexperienced

Student Running

Tests

Lack of knowledge or

experience by test personnel

resulting in damage to lab

equipment, facilities,

components of the launch

vehicle, or other team members

or students

3D

All tests have been done by

supervisors of each group.

Multiple team members will be

present during testing in order to

ensure proper protocols are

followed and safety precautions

are taken

Subsection team leads will be

presents thought the duration of

their area's testing to ensure

proper usage of lab equipment.

4D

Wind Tunnel

Operation at

Excessive Speeds

Lack of knowledge or

experience by test personnel

resulting in damage to the testing

apparatus or components of the

rocket being tested in the wind

tunnel

4D

All tests have been done by

supervisors of each group.

Multiple team members will be

present during testing in order to

ensure proper protocols are

followed and safety precautions

are taken

Aerodynamics subsection team

leads will be presents thought

the duration of wind tunnel

testing to ensure proper usage of

lab equipment in accordance

with manufacturer specifications

4E

Debris in the Wind

Tunnel

Failure to properly clean and

inspect wind tunnel prior to

testing resulting in potential

damage to testing apparatus,

components of the rocket or

harm of individuals running the

test

4B

Proper PPE and eye protection

must be worn at all times in the

lab. In the event that debris does

fly out of the wind tunnel during

testing, multiple students will be

present to assist in the clean-up

of the debris.

Safety officer will monitor

testing to ensure proper PPE,

such as gloves, ear and eye

protection are being worn during

testing operations.

4D

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Overuse of Wind

Tunnel

Excessive testing beyond

apparatus capabilities causing

damage to the testing apparatus

that could result in inability to

conduct future tests

3D

Wind tunnel testing will be

scheduled in advance with

breaks in between tests to allow

the engine to properly cool. The

tunnel will also be inspected

before testing to ensure proper

conditions.

Aerodynamics subsection team

leads will be presents thought

the duration of wind tunnel

testing to ensure proper usage of

lab equipment in accordance

with manufacturer specifications

3E

Black Powder Fails

To Ignite

Improper storage of black

powder leading to no separation

or deployment of parachute, thus

creating a potential for

catastrophic damage to launch

vehicle or injury to spectators

2D

Secondary charges can be used

in order to ensure that if one

change fails another can engage

to deploy the parachutes.

See ejection testing subsection

of FRR regarding proper black

powder handing

3D

Excess of Black

Powder Used in

Testing

Failure to properly measure

correct amount of black powder

for sample resulting in full

separation of rocket leading to

damage to various components

and potential failure of other

systems and debris

2D

Manufacturer specification are

followed in order to determine

how much black powder is need

to pressurize the rocket based on

its weight

See ejection testing subsection

of FRR detailing the amount of

black powder to be used for

complete and optimal separation

2E

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Failure to Properly

Secure Payload

Improper connection inspection

prior to launch leading to

potential damage to the payload

or its housing container. Could

compromise the structural

integrity of various components

or lead to failure of other

operations

3C

Prior to testing the pre-launch

checklist were used to verify that

payload is mounted correctly in

place and all connections are

secure to ensure safe launch

operations

In accordance with launch

procedures checklist, the

payload will be inspected prior

to launch for secure attachment

and flight readiness

3D

Parachute is Not

Packed Properly for

Testing

Incorrect packing procedure or

method used resulting in failure

in parachute to deploy at proper

altitude resulting in launch

vehicle experiencing more force

than planned upon landing

1C

Multiple tests have been

conducted in order to verify the

packing method and give the

student practice with packing the

parachute into the body tube. On

launch day, the pre-launch

checklist will be followed to

ensure proper packing of the

parachute in accordance with

standard practices.

Testing have been done in order

to verify the packing method and

give the student practice with

packing the parachute into the

body tube. On launch day, the

pre-launch checklist will be

followed to ensure proper

packing of the parachute in

accordance with standard

practices.

3D

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Table 22 - Failure Modes and Effects Analysis - Launch Support Equipment

Risk/Hazard Root Cause/Effect

Severity/

Likelihood

Proposed Mitigation Verification Plan Post-Control

Severity/Likelihood

Instability in Guide

Rail

Faulty component(s) or failure to

account for various changes to

rocket made between launches

can cause launch vehicle to

deviate from its deal path

potentially leading to

endangerment of spectators or

damage to components

2C

Prior to launch operations, the

guide rails will be inspected by

the lead safety officer in

accordance with the pre-launch

checklist in order to ensure safe

operations

In accordance with set up on

launch pad checklist, the launch

pad and guide rail will be

inspected for structural flaws or

bowing that could lead to

instability in launch and

deviation of the rocket from its

ideal flight path

3D

Improper Transport

of Launch Vehicle

Lack of care when handing

launch vehicle or components

resulting in potential damage to

launch vehicle and/or

compromise of structural

integrity of individual

components

2C

The launch vehicle will be

transported in its specially made

container which will provide

support for all fragile areas of

the rocket, while protecting it

from slipping, vibrations or other

potentially damaging impacts.

In accordance with final

assembly checklist, prior to

launch, all subsection will be

inspected by their team lead for

cracks, chipping, or other

structural flaws that could have

been acquired during transport

3D

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Improper Storage of

Launch Vehicle

Lack of care and safety when

storing launch vehicle or

components leading to damage

to launch vehicle and potential

compromise of structural

integrity of components

2D

The launch vehicle is stored in a

specific container in a locked

room during fabrication and in

between tests. Only team leads

and safety officers have access

to the room to prevent the rocket

from being mishandled. The

room is kept at room

temperature to not adversely

affect any components

In accordance with the parts

checklist, the safety officer will

periodically inspect storage

cabinets as well as launch

vehicle holders to ensure all

supplies have been returned

following use, and are being

stored in their proper place

3D

Improper Transport

of Rocket Motor

Handling of motor not in

accordance with specifications

leading to potential damage to

payload or other essential

components upon launch

3C

The rocket motors will be

transported in a fireproof case

that will prevent moisture for

getting into the motor. The case

will also protect the motors

against slipping, vibrations, and

other potentially damaging

impacts.

In according with motor

preparation checklist, the rocket

motor will be stored off site and

will be transported to the launch

location in a protective,

waterproof case

3E

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Improper Storage of

Rocket Motor

Potential damage to payload or

other essential components upon

launch due to improper

placement of motor in

potentially compromising

locations

2D

The rocket motors are stored in a

fireproof case. This case is kept

in a locked cabinet in order to

prevent from other team

members handing the motors.

This cabinet will remain at room

temperature and dry in order to

not allow heat or moisture

adversely affect the motor.

In according with motor

preparation checklist, the rocket

motor will be stored off site and

will be transported to the launch

location in a protective,

waterproof case

3E

Improper Handling

of Rocket on Launch

Pad

Handling of launch vehicle not

in accordance with guidelines

listed on pre-launch checklist

leading to endangerment of

spectators, and minor to

catastrophic failure of the rocket

and its subsystems

2D

Only trained and essential team

members will handle the rocket

during launch operations. Pre-

launch safety checklists will be

used in order to ensure everting

is safe for launch

In accordance with motor

preparation checklist, the leader

of the propulsion subsection will

retrieve the rocket motor from its

protective, waterproof casing,

and will ready the motor for

ignition after all subsequent

subsection inspections have been

completed

2E

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Instability of Launch

Pad

Faulty component or failure to

properly inspect launch pad prior

to flight leading to potential for

rocket to deviate from ideal

flight path endangering

spectators and causing failure of

components or drift

2B

Prior to launch operations, the

launch pad will be inspected by

the lead safety officer in

accordance with the pre-launch

checklist in order to ensure safe

operations

In accordance with the set up on

launch pad checklist, prior to

launch operations, the launch

pad will be retrieved from

NASA will be inspected by the

lead safety officer in conjunction

with the RSO for any structural

flaws that could lead to

instability in launch operations

2C

Faulty Ignitor Clips

Improper handling or storage of

component causing rocket to be

unable to ascend off of launch

pad

3C

Prior to launch operations, the

ignitor clips will be inspected by

the lead safety officer in

accordance with the pre-launch

checklist in order to ensure safe

operations and successful launch

In accordance with ignition

checklist, ignitor clips will be

inspected prior to attachment to

the launch vehicle by propulsion

subsection lead.

3D

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Table 23 - Failure Modes and Effects Analysis - Launch Operations

Risk/Hazard Root Cause/Effect

Severity/

Likelihood

Proposed Mitigation Verification Plan

Post-Control

Severity/Likelihood

Cracking in Main

Body Tube

Improper storage of launch

vehicle or transportation leading

to compromise in the structural

integrity of the rocket leading to

potential damage to other

components or failure of other

subsystems

2D

Prior to launch, the body tube

will be thoroughly inspected for

cracking, splintering, or

fatiguing in according with the

procedures listed in the pre-

launch checklist in order to

ensure safe launch operations

In accordance with the launch

procedures checklist, the main

body tube will be inspected by

the safety officer prior to launch

for any structural imperfections

3D

Gaps Between

Connecting Pieces

Failure to fabricate subsections

according to specifications

yielding an inability to assemble

components of rocket properly

with secure attachment

potentially leading to failure of

rocket upon launch or testing

operation.

2C

All components of the rocket

were measured after fabrication

in order to ensure they met the

dimensions specified in the

design. The pre-launch safety

checklist will be used to ensure

team members visually inspect

connections of components prior

to launch

All components of the rocket

have been measure after

fabrication in order to ensure

they meet the dimensions

specified in the design. The pre-

launch safety checklist will be

used to ensure team members

visually inspect connections of

components prior to launch

3C

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Collision with

Object in Sky (Tree,

Bird, Etc.)

Failure for parachutes to deploy

at proper altitude resulting in

damage to launch vehicle and

compromise in structural

integrity of impacted

components

2C

Test have been run in open areas

to ensure no overhanging trees,

roofs, or other items could

impede the flight operations.

Furthermore, subsequent tests

have been completed in order to

determine the strength of the

body tube and nosecone so that

in the event the launch vehicle

does strike a bird it can

withstand impact and return

safely

See scale model testing

subsection of FRR report

regarding impact testing of

launch vehicle

2D

Instability During

Flight

Failure of team members to

account for ballast adjustments

made prior to launch resulting in

change of center of gravity and

leading to inability of the rocket

to maintain its projected flight

path

1B

Maintain safe distance from

launch pad

In accordance with post-flight

inspection, all team members

will wait until rocket has landed

in a safe location before leaving

safe launch zone

2C

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Altimeter or Other

Electronics in

Avionics Bay

Malfunction/Fall Off

Failure to properly inspect

security of attachment between

electronics, or test functionality

prior to launch causing potential

short circuiting or harm to

spectators below

3B

Verify all electronics work

properly before launch and are

firmly attached to the rocket

In accordance with launch

procedures checklist, all

electronics will be tested for

functionality prior to launch

operations. For further detail on

electronics testing, see altimeter

testing section of FRR

3C

Coupler Excessively

Tight

Parts not fabricated according to

specifications resulting in

potential failure of parachute to

deploy leading to damage to

rocket

2D

Run multiple tests to ensure

proper amounts of black powder

is used to allow rocket to

separate

In accordance with launch

procedures checklist, the ability

of adjoining sections to separate

will be tested. For further detail

on proper ejection charges for

separation, see ejection testing

subsection of FRR

3D

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Environmental Considerations

Additionally, when considering the safety and impact of the rocket, considerations must be

given to how the vehicle will impact the environment, and how the environment will impact the

vehicle. These considerations are represented below in the environmental hazard analysis, shown

in Table 24.

Table 24 - Environmental Consideration Hazard Analysis

Risk/Hazard Root Cause/

Effect

Severity/

Likelihood

Mitigation and

Control

Post-Control

Severity/Likelihood

Vehicle Effects on Environment

Epoxy Fumes

Fumes released

during

construction

resulting in

hazardous working

conditions created

for team members

as a result of toxic

air

4A

Work in well

ventilated spaces

and dispose of

waste properly

4D

Epoxy Not

Disposed of

Properly

Failure to follow

proper disposal

protocols leading

to potential fire

hazard and damage

to lab or

equipment

4C

Let epoxy fully

cure before

disposal in order

to prevent fire

hazard

4E

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Dust Particles

Fabrication

operations

producing small

dust particles from

sanding or

machining

operations are

released into the

environment which

can result in

breathing problems

4A

Wear mask when

sanding to avoid

inhaling dust

particles and try

to contain dust

when sanding

opposed to freely

releasing it into

surrounding air.

4D

Rocket Motor

Ignition

Failure to properly

secure motor,

therefore, upon

ignition, when

motor reaches high

temperatures and

hot exhaust is

released, the motor

could become

displaced or burn

the areas where the

rocket is launched

or lands

2D

Place flame

resistant material

beneath the

launch pad to

avoid burning the

immediate

surroundings or

starting a fire

3D

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Debris from

Rocket

Component failure

leading to

fragments of the

rocket breaking off

during flight or

upon landing

impact and

becoming

irretrievable,

leading to minor

environmental

harm due to

inability to

decompose and

toxicity of

component

3D

Ensure fully

functioning

parachutes before

launch via pre-

launch recovery

preparation

checklist and

check to make

sure all

components of

the rocket and

payload are

accounted for

upon return in

accordance with

post-flight

inspection

checklist

3E

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Environmental Effects on Vehicle

Water

Improper storage

or launching in

unfit conditions

can lead to water

exposure, which

can cause

malfunctioning of

electronics within

the avionics bay,

or damage to the

body of the rocket,

which will be

constructed out of

Blue Tube that is

not 100% water

resistant

2E

Avoid launching

rocket in wet

conditions and

store rocket in

proper stand in a

dry area for

storage and

transport

3E

Wind

Launch into

excessive wind

speeds leading to

deviation from

launch vehicle's

ideal flight path

thus leading to

damage to the

rocket and

potential harm to

spectators

3C

Avoid launching

rocket on days of

high speed winds

or unpredictable,

strong wind gusts

3E

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Humidity/Moisture

Improper storage

of components

resulting in

potential corrosion

and weakening of

various materials

used to construct

the rocket. This

moisture can also

negatively impact

on-board

electronics

3D

Store rocket in a

dry area to avoid

moisture entering

the rocket over

time via humid

air

4E

Visibility

Launch during

times of low cloud

coverage resulting

in inability to track

the rocket thus

leading to debris

not being retrieved

and damaging the

environment

4C

Avoid launching

rocket on days

with low cloud

coverage and

poor visibility

4E

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General Risk Assessment

Finally, a general risk assessment, shown in Table 25, was conducted in order to account for

various extraneous risks not accounted for in previous sections, such as time, resources, the

budget, scope, and functionality.

Table 25 - General Risk Assessment

Risk/Hazard Root Cause/

Effect

Severity/

Likelihood

Mitigation and

Control

Post-Control

Severity/Likelihood

Limited

Resources

Being a first year

team with a small

budget could lead

to a lack of quality

design or

fabrication

material and to

failure to meet

mission objective

or overall poor

performance

2C

The team has work

with faculty members

as well as local

rocketry club members

in order to gain a better

understanding of

rocketry and develop a

functional rocket.

2C

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Tight or

Minimal Budget

Being a first year

team with minimal

established

funding, the team

could be forced to

use parts that are

not optimal, or be

unable to replace

parts of the rocket

that are broken

during testing

3A

The team and its adult

educators have applied

for and been given

grants in order to fund

parts of the project.

Additionally, the team

has held fundraisers to

provide the team with a

flexible budget beyond

the normal amount of

money allotted to the

project by the school

3B

Mismanagement

of Time

Inability to

manage project on

a weekly basis

could potentially

lead to major

delays resulting in

the quality of work

lacking, or the

rocket not being

completed by

competition

1E

Team members have

and will continue to fill

out weekly time cards

and log their hours in

the task breakdown in

order to ensure

everyone remains on

schedule

2E

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Underestimation

of Scope of

Work

Inability to budget

time properly

could leave the

project running

behind schedule

and various facets

of the rocket not

being completed in

a quality manner

2E

There has been and will

continue to be constant

communication

amongst all team

members and with

NASA project leads to

ensure the scope of

work is clear and

everyone stays on task

3E

Increase in

Safety

Regulations

Failure to meet

proper FAA and

NASA safety

regulations could

lead to team to be

forced to add

material to the

rocket in order to

increase safety,

which will result

in an increase in

expenses

2D

The team has designed

and downselected with

safety as the foremost

priority, and will

clearly identify all

safety measures before

all operations so that

additional, last-minute

safety measures do not

have to be taken that

will inflate the budget.

2E

105 | P a g e

Launch Operations Procedures

Parts Checklist

In order to ensure safe and uninterrupted transportation of components and launch day

procedures, parts checklists were developed for each of the following subsections: propulsion,

aerodynamics, main payload, avionics bay/electronics payload, recovery, and safety and

education, as well as a miscellaneous checklist to account for various extraneous items that the

team will need for launch day operations. These checklists, which can be seen below in Table 26

through Table 32, were developed by each subsections respective team lead in conjunction with

the safety officer to ensure all vital parts of the launch vehicle, as well as supporting materials,

are accounted for and available for use on launch day.

Table 26 - Parts Checklist - Propulsion

Initial Part Quantity

Liner for Motor Case 1

Motor Case 1

Aft Closure 1

Bow Closure 1

Forward Seal Disc 1

Reload Kit 1

Grains 3

Ignitor 2

Retention System Cap 1

106 | P a g e

Initial Part Quantity

Water -

Rags 3

Pocket Knife 1

Flat head screw Driver 1

Super Lube Synthetic

Grease

2

Wire Strippers 1

Box Cutter 1

Table 27 - Parts Checklist - Aerodynamics

Initial Part Quantity

3/36" Hex Key 1

3/16" in Hex Bolts 6

Nose Cone 1

Bow Body Tube 1

Aft Body Tube 1

Body Tube Holders 2

Spare Fin 1

JB Weld Tube 2

Extra Rail Buttons 4

Launch Rail 1

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Table 28 - Parts Checklist - Main Payload

Initial Part Quantity

CR1-400 Wire Rope

Isolators

12

5.36" Blue Tube (Cylinder

2)

1

Base Springs (#866) 5

Spacer (clear acrylic) 1

Recovery Bolts 3/8" x 1.25"

Length

2

Spring Fastening Bolts For

3/8" x 16 Bolt, 1/4" Height

30

Spring Fastening Washers 30

Bulkheads 2

Pins 0.1405 inch diameter 24

Aluminum Squares 1x1x0.1

inches

12

3D Printed Cylinder

(Cylinder 1)

1

3D Printed Cap 1

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Table 29 - Parts Checklist – Electronics Payload/Avionics Bay

Initial Part Quantity

Atlus Metrum TeleMega 1

Starter Pack 1

Arrow 440-3 Yagi Antenna 1

SMA to BNC Adapter 1

10-24 9/16" O-Ring Bolts 4

5-40 5/8" Altimeter Bolts 4

O-Ring 1

1" Long, 0.25x40" Studs for

Ballast

4

109 | P a g e

Table 30 - Parts Checklist - Recovery

Initial Part Quantity

Coupling Tube 1

Electronics Sled 1

Ejection Charge Igniter 8

Plastic Bags 4

Flat Washers 4

Lock Washers 2

Wing Nuts 2

1/4" Hex Nut 1

Shear Pins 12

1/4" Quick Links 6

35' Recovery Harness 2

Nomex Sleeves 2

Nomex Squares 2

24" Drogue Parachute 1

96" Main Parachute 1

Roll of Masking Tape 2

Black Powder Assurance Recovery Fiber Sheets 25

110 | P a g e

Table 31 - Parts Checklist - Safety and Education

Initial Part Quantity

First Aid Kit 1

Fire Blanket 1

Fire Extinguisher 1

Safety Glasses 15

Ear Plugs 15

Dust Mask 5

Table 32 - Parts Checklist - Miscellaneous

Initial Part Quantity

Folding Table 1

Chairs 4

Quick Dry Epoxy Tub 2

111 | P a g e

Final Assembly Checklist

Below, in Table 33 through Table 40, are final assembly checklists for each subsection that

were used for full-scale rocket assembly prior to launch to ensure safe and successful operations.

For each checklist, the leader of the subsection is required complete each check-off point, in the

order that they appear on the list, and then present the list to the safety officer for approval and

sign-off. After this, the next checklist can be completed. It is important to note that each

checklist is to be completed one at a time, in the order that they appear in this document, and not

in parallel with other checklists currently in progress. In the event that any point on the checklist

cannot be completed, the subsection team lead should immediately notify the safety officer so

that the problem can be dealt with according to the procedures listed in the troubleshooting tables

(Table 46 through Table 49). After all pre-launch checklists and inspections have been

completed and approved by the safety officer, launch operations may commence.

Table 33 - Final Assembly Checklist - General Set Up

Initial Check-Off Point

Set up table for launch vehicle preparation and pre-launch inspection

Equip all personnel handling the launch vehicle with proper PPE equipment

Inspect all members for safety glasses, gloves, and proper attire before handling

any launch vehicle-related supplies

Unpack all supplies and boxes from the truck

Separate supplies by subsection

Remove launch vehicle from transport case and transport to housing on

inspection table

112 | P a g e

Table 34 - Final Assembly Checklist - Comprehensive Structural Inspection

Initial Check-Off Point

Visually inspect body tube for cracks, bumps, abrasions or any other

imperfections that could have been acquired during transport that could

adversely affect the flight of the rocket

Physically inspect rocket tube for structural integrity and flight readiness

Inspect fins for any structural imperfections or bowing that could have been

acquired during transport

Physically inspect nosecone for cracks, chipping, or any other damage that

could have been acquired during transport and handling

Examine thrust plate and couplers for solid connection and structural integrity

Table 35 - Final Assembly Checklist - Electronics

Initial Check-Off Point

Inspect avionics bay for flaws or damage to ensure nothing was broken or

disconnected during transport

Ensure proper connection of all electrical wires by inspection and comparison

to wiring diagrams

Test avionics for proper functioning

Assemble avionics bay and check for proper connection to shock cord

Test GPS tracking device and altimeter to ensure proper functioning

Secure avionics bay using proper fasteners

113 | P a g e

Table 36 - Final Assembly Checklist - Payload

Initial Check-Off Point

Examine payload housing container for any structural imperfection that could

have been acquired during transport

Inspect wire rope isolators for fraying or fatigue

Visually examine springs on payload housing container for structural integrity

Ensure proper filling of dampening material to protect payload

Check for secure connection between fragile material protection apparatus and

recovery section

Table 37 - Final Assembly Checklist - Recovery System

Initial Check-Off Point

Inspect drogue parachute and shock cord for any imperfections or tears that

could lead to error in recovery operations

Examine connection between drogue parachute shock cord and main body

section

Examine connection between drogue parachute shock cord and drogue

parachute

Fold and pack the drogue parachute

114 | P a g e

Initial Check-Off Point

Wind excess drogue parachute shock cord to ensure proper deployment of

drogue parachute

Inspect main parachute and shock cord for any imperfections or tears that could

lead to error in recovery operations

Examine connection between main parachute shock cord to main body section

Examine connection between main parachute shock cord and main parachute

Fold and pack the main parachute

Wind excess main parachute shock cord to ensure proper deployment of main

parachute

115 | P a g e

Table 38 - Final Assembly Checklist - Motor/Ejection System Preparation

Initial Check-Off Point

Inspect individuals preparing motor for proper PPE, including glasses, gloves,

and mask

Remove black power container from storage case

Check black powder to ensure no moisture has compromised the sample

Measure and pour 2 grams of black powder into charge cup to be used for

drogue parachute

Measure and pour 3 grams of black powder into charge cup to be used for main

parachute

Inspect motor casing for any structural imperfection acquired during transport

Remove motor from storage container

Examine motor and casing to ensure it is not wet or containing any moisture

that could cause misfire or deviation from ideal flight path

Assemble motor following manufacturer specifications

Install motor into launch vehicle

116 | P a g e

Table 39 - Final Assembly Checklist - Secure Attachment Inspection

Initial Check-Off Point

Check for secure attachment between motor and casing

Examine nosecone for level and secure attachment with main body tube

Inspect electronics bay within nose cone for proper fastening

Inspect for proper connection between nosecone and payload bay

Check for secure attachment between main payload and recovery system

Inspect all exterior connections and assemblies on the rocket for proper fitting

Table 40 - Final Assembly Checklist - Launch Pad/Pre-Launch Inspection

Initial Check-Off Point

Transport launch vehicle to Range Safety Officer for inspection

Continuity test igniter clips for proper functioning with launch controller

Inspect launch rail for bowing or imperfection that could cause the rocket to

launch in an unplanned direction

Connect the ignitor clips to the motor ignitor

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Motor Preparation

In order to prepare the motor for ignition and launch operations, the following checklist,

shown below in Table 41, was used.

Table 41 - Motor Preparation Checklist

Initial Check-Off Point

Remove motor from protective, waterproof casing

Assemble motor according to manufacturer specifications

Remove the top of the screw on the retention system

Place motor into inner tube with the nozzle facing the rear of the rocket in the

open-air

Examine placement of motor in inner tube to ensure secure fit

Screw top of retention system back into place

Place cap around nozzle to ensure a moisture does not enter the grains

Motor is ready for ignition

118 | P a g e

Recovery Preparation

To prepare the recovery system for launch operations the recovery preparation list, displayed

below in Table 42, was used.

Table 42 - Recovery Preparation Checklist

Initial Check-Off Point

Test each battery with a multimeter to ensure that it is fully charged to 9 volts

Reconnect each battery to its respective altimeter

Insert the mounting sled into the coupling tube by sliding it over the threaded

steel rods

Connect mating female molex plugs with their male counterparts from the

altimeters

Electrical connections for the drogue and main ejection charges are established

Attach aluminum bulkhead with lock washer and wing nuts

Assemble the coupling tube

Open end of coupling tube is now sealed

Measure two 2.00 g black powder samples to be used for the drogue charges

Place sample into small plastic bag with an ignitor

Measure two 3.00 g black powder samples to be used for the main charges

Place sample into small plastic back with an ignitor

Twist each bag to compress the black powder around the tip of the ignitor

Insert each ejection charge into ejection well

Insert foam insulating material to hold each charge in place

119 | P a g e

Initial Check-Off Point

Seal each ejection well using masking tape

Strip electrical leads

Clamp electrical leads to terminal block

Attach recovery harnesses

Secure quick links on the end of each harness to U-bolts on the body and

coupling tubes

Wrap each harness in a spiral form

Insert the wrapped harness into the body tube

Wrap main parachute in Nomex flameproof fabric

Insert main parachute into launch vehicle

Wrap drogue parachute in Nomex flameproof fabric

Insert drogue parachute into launch vehicle

Insert the coupling tube into the aft body tube

Secure the coupling tube and aft body tube using two nylon shear pins

Fit aft body tube onto top of the coupling tube

Secure the aft body tube and coupling tube using two nylon shear pins

Activate altimeter alarming switches through exterior holes in the coupling tube

120 | P a g e

Setup on Launch Pad

After all subsections of the rocket had been properly configured, Table 43 was used in order

to ensure proper safety procedures were followed when transporting the launch vehicle to the

launch pad and when preparing the rocket for launch operations.

Table 43 - Launch Pad Configuration Checklist

Initial Check-Off Point

Obtain launch pad for official competition from NASA

Set launch box down at safe viewing distance

Inspect the launch rail for any structural flaws that could cause the rocket to

deviate from its ideal course of travel

Lower launch rail height for safe rocket insertion

Transport launch vehicle to launch pad with approved team members

Place the launch vehicle on the launch rail

Insert launch rail onto base of launch pad

Secure launch rail to base of launch pad with two threaded bolts

Adjust launch pad to vertical setting using the design feature on the base of the

launch pad

All non-level two members retreat to safe launch zone

Complete ignitor installation checklist

Arm rocket for launch

Remaining members retreat to safe launch zone

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Ignitor Installation

After the launch vehicle was properly configured on the launch pad and non-level two

members of the team had retreated to the safe viewing area, the ignitor was installed in

accordance with the checklist in Table 44.Table 44 - Ignitor Installation Checklist

Table 44 - Ignitor Installation Checklist

Initial Check-Off Point

Strip ignitor wires 2 inches to allow for more surface contact with the

composite for ignition

Remove paper around the end of the ignitor from the composite

Insert ignitor into motor

Inspect ignitor to ensure entire ignitor is within the grains of the motor

Pinch ignitor wires where end of the wires reach the end of the motor

Remove pinched wire from the motor

Measure pinched wire length

Check to ensure that pinched wire length is the matches up with the length of

grains in the motor

Replace measured wire back into motor

Attach stripped wires to ignition system

Wrap stripped part of the wires around the system to allow for proper surface

contact

Inspect continuity of system

Connect ignitor leads to launch controller

122 | P a g e

Initial Check-Off Point

Ignitor and ignition system is set-up and ready for launch

123 | P a g e

Launch Procedures

Following the installation of the ignitor, the rocket was armed and ready for launch. The

launch procedures checklist, below in Table 45, contains all of the necessary checkpoints that

must be met in order to ensure a safe and successful launch. To ensure the safety of all team

members as well as spectators, equipment, and facilities, all check-off points listed in the final

assembly checklist and launch procedures checklist must be initialed by subsection leaders in

order for launch operations to commence.

Table 45 - Launch Procedures Checklist

Initial Check-Off Point

Ensure a safe working area before transporting rocket to the launch pad

Check the safety and readiness of team members and bystanders by ensuring

proper PPE and safety glasses are worn by all individuals transporting the

rocket

Carefully transport rocket to launch pad

Visually inspect the rocket main body tube for any structural imperfections

Visually inspect the fins for any structural imperfections

Inspect launch vehicle for proper connections between all sections of the rocket

Test nosecone and body tube's ability to separate

Examine main body tube for flight readiness

Inspect fins for flight readiness

Inspect nosecone for flight readiness

Review payload to ensure flight readiness

124 | P a g e

Initial Check-Off Point

Test electronics (GPS, camera, altimeter, etc.) to ensure they are armed and

functional prior to launch

Inspect launch pad and guide rails for readiness

Place rocket on launch pad

Have non-level two team members move away from the launch pad back to the

safe-viewing area

Arm the rocket motor for ignition

Disarm all safeties on the rocket

Have remaining team member retreat to safe-viewing distance to watch launch

Check with Range Safety Office to ensure all codes and rules ae met and the

rocket is clear for launch

Initiate rocket ignition

Check for proper ignition

Watch flight so that launch vehicle sections do not get lost

Recover payload and main body section after landing

Disarm altimeter and any unfired charges

Disassemble launch vehicle

Inspect launch vehicle for any cracks, breaks or fatigue as a result of testing

Record altimeter data

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Troubleshooting

Table 45-Table 49 below, detail troubleshooting tactics that can be used to address common

problems that could be encountered during the pre and post launch subsection inspections.

Table 46 - Troubleshooting - Cracking in Main Body Tube or Subsection

Initial Check-Off Point

Replace cracked part if spare part is available

Evaluate severity of structural compromise

Determine if cracked piece is load bearing

If not load bearing, epoxy part

If cracked part is critical and load bearing, postpone launch until replacement

part can be obtained or manufactured

Table 47 - Troubleshooting - Insecure Fit Between Adjoining Subsections

Initial Check-Off Point

If too large, sand oversized subsection down until secure fit is reached

If too small, replace with spare part

If spare part is unavailable and part is too small, add layers or tape to increase

diameter until secure fit is reached

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Table 48 - Troubleshooting - Unresponsive or Malfunctioning Electronics

Initial Check-Off Point

Inspect wiring to see if there is any disconnect or break in the circuit

Test battery to ensure it is operating at the proper voltage

Inspect wiring switch

Examine wiring terminals for crossed wires or insertion into incorrect ports

Replace unresponsive/malfunctioning electronic piece

Table 49 - Troubleshooting - Insecure Connection Between Launch Rail and Launch Pad

Initial Check-Off Point

Inspect launch pad for debris that could be limiting proper connection

Inspect launch rail for bowing that could be limiting proper connection

Screw threaded bolts further into launch pad to create more secure connection

If connection is still not secure, drill new holes to screw threaded bolts into

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Post-Flight Inspection

Following flight operations and retrieval of the rocket, all areas of the rocket will be inspected in

order to determine the success of the team’s testing and design, as well as individual component

suitability to be reused on a subsequent flight. In order to complete this post-flight inspection,

Table 50 is used.

Table 50 - Post-Flight Inspection Checklist

Initial Check-Off Point

Wait until rocket has landed in a safe location before leaving safe launch zone

If the rocket is not deemed safe for retrieval by RSO, stay in safe launch zone

and have proper individuals retrieve rocket

If the rocket is deemed safe for retrieval by RSO, have the safety officer

approach launch vehicle for retrieval

Retrieve launch vehicle and return to working area for inspection

Remove motor casing once it reaches a temperature that is cool enough to

handle

Inspect motor casing for cracking or other structural flaws

Clean motor casing

Disassemble the rocket into individual subsection

Remove altimeter from the rocket

Record the official altitude of the launch vehicle following flight operations as

measured by the altimeter

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Initial Check-Off Point

Aerodynamics team inspects main body tube, fins, and couplers for cracking or

structural flaws acquired during flight

Main payload team inspects payload for structural integrity and security of

fragile material

Electronics payload team inspects altimeter and avionics bay for proper

functioning and any damage to electronic systems as a result of flight

operations

Recovery team inspects all components of the recovery subsection

Safety officer completes overall inspection of all subsection inspections

Receive all good from RSO

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Project Plan

Testing

The testing plan outlined in the CDR has almost been completed. All but two tests have been

completed, and each can be seen in greater detail below. Table 51 summarizes each test and its

results.

Table 51 - Test Results

Test Data Taken Status and Results

Altimeter Testing Altitude, GPS tracking,

and live feed

All three altimeters

precisely measured altitude.

The GPS tracking and live

feed worked properly.

Complete.

MTS Bulkhead Testing The force required for

failure of the assembly.

The epoxy failed, not the

carbon fiber or aluminum.

Complete.

Ejection Testing

If separation is achieved,

the amount of black powder

needed.

Separation was achieved

10 times for each body tube.

Complete.

Parachute Force

Deployment Testing

Force of the parachute

deployment. In Progress.

Wind Tunnel Testing Strain from a strain

gauge. Incomplete.

Scale Model Testing Full System Test Two Successful Flights.

Complete.

Payload Spring Testing Spring Constant Check Complete spring constant

matches. Complete.

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Test Data Taken Status and Results

Full Scale Testing Full System Test Three Successful Flights.

Complete.

Payload impact testing Modified Charpy Impact

Test Complete

Altimeter

The main scoring altimeter and the recovery altimeters were re-tested with the drone as they

were for the subscale launch. The process for this test can be found in the CDR. Attaching to the

drone allowed all three altimeters to be tested to ensure the GPS, altitude reading, and live feed

all worked correctly. The GPS tracking and live feed is only on the main scoring altimeter and

both worked correctly.

The altitude data from the drone flight averaged two feet higher than the three altimeters,

however the altimeters measured the same altitude. The difference in height makes sense because

the altimeters were suspended below the drone two to three feet depending on which test number

it was. The drone test was repeated five times with the lengths of the rope being measured after

the altimeters were tied off.

Along with the altitude being tested, the flight data from the recovery altimeters also showed

where the parachutes would have been deployed. The deployment altitudes demonstrates that not

only is the altimeter reading altitude, but both recovery altimeters have been correctly set to

deploy at the proper altitude.

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MTS (Bulkhead)

Tensile testing with the MTS Machine was designed to determine, which component would

fail and how much force is required to cause failure. Knowing how much force will cause failure

will verify the manufacturers’ specifications. The test also shows that nothing should fail in

flight because all components have been designed and selected to withstand more stress than

what will be endured in flight.

The assembly was manufactured with a small piece of the body tube, two spare recovery

bulkheads, and two identical U bolts. The two bulkheads were epoxied into the body tubes, with

one at each end. These bulkheads are identical to what will be used in the full-scale flight. The U

bolts were attached to the bulkheads in the same manner as the full scale.

Two pieces of fracture mechanics clevis grip were used to mount the U bolts in the MTS

machine. Figure 40 shows the clevis grips attached to the U bolts before mounting into the MTS

machine. Figure 41 shows the assembly mounted into the MTS machine. The MTS test was

repeated twice, on two identical assemblies. To ensure the data was as consistent as possible the

angle of the bulkhead was measured while it was attached to the MTS machine, the first angle

measured 7.4 degrees and the second measured 7.9 degrees. With both tests the epoxy failed

first, which is called adhesive failure. The test is considered to be successful because the epoxy

failed first and at a force greater than what it will endure in flight.

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Figure 40 - Fracture mechanics clevis grip attached to U bolts Bulkhead assembly for MTS testing

133 | P a g e

Figure 41 - The Assembly Mounted into the MTS Machine

The test was determined to be successful if the failure is a higher force than what will

be experienced during flight. If the MTS machine reached the maximum travel distance, and the

assembly did not fail, then the maximum force put onto the assembly would determine if the test

was a success. Table 52 shows the results from the MTS test. The OpenRocket simulation

showed that the parachute ejection should put a force of 400 lbf onto the rocket body. Using data

from the full scale flight, the actual force felt on the rocket was 206 lbf. With both MTS tests, the

bulkheads withstood a significantly higher force than what will be experienced in flight.

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Table 52 - MTS Test Results

Maximum Force 2252.838 lbf 1555.44 lbf

Component Failure Epoxy Epoxy

The bulkheads were tested in order to make sure that extra inspections are done at the point

of failure before and after the flight. Safety is the primary consideration and locating the most

likely point of failure allows the team to ensure safe flights. The maximum force for the first test

was 1555.436 lbf, and the maximum force for the second test was 2252.838 lbf. Both assemblies

used were made from the same materials, however keeping the exact same amount of epoxy is

impossible. On the second test more epoxy was used to better represent the actual amount on the

recovery bulkhead in the rocket. On the second test before epoxying the bulkhead into the carbon

fiber, both pieces were roughed up with a file. Roughing up each piece allows the epoxy to

adhere better compared to two smooth surfaces. The increase in epoxy, along with the rougher

surface area, is what caused the higher force needed to fracture the assembly.

The procedure used to run the MTS Machine and perform the tensile test on the assembly can

be found in Appendix K.

Ejection Testing

To be sure that the entire recovery system would function as designed, multiple ground

ejection tests were performed for each body tube and altimeter. A successful ejection test

consists of complete ignition of the black powder charge and separation of the body tube from

the coupling tube. Satisfactory performance of each altimeter signal is attained through two

135 | P a g e

successful tests. These tests ensure that the wiring of the recovery electronics is sound and that

the parachute compartments are sufficiently airtight. Additionally, the test will test the shearing

of the nylon pins holding the body tubes together.

The size of the ejection charges were determined using equations available through the

website of the Nevada Aerospace Science Association. Based upon the guidance of the NAR

members at Mid-South Rocket Society, the mass of black powder used in each charge will be

double the calculated mass.

Before the test can begin, an ejection charge must be packed according to the procedure

described in the Recovery Preparation section. After inserting the coupling tube into the body

tube to be tested, the coupling tube was braced between two sandbags. This ensured that the

coupling tube remained stationery during the test, preventing damage to the electronics. The

body tube was rested on an adjacent sandbag. To slow the body tube after ejection and protect it

from external damage, a series of cloth dampers were hung in the path of the body tube’s motion

away from the coupling tube. The USB data transfer kit was connected to the altimeter and a test

signal was fired.

The ejection tests were entirely successful with the exception of a single backup circuit test

of the drogue parachute charge, which led to the re-soldering of a disconnected wire. An

overview of the tests is given in Table 53.

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Table 53 - Results of ejection testing

Signal Number of

Tests

Number of

Failures Notes

Primary Main 2 0

Primary Drogue 2 0

Backup Main 2 0

Backup Drogue 3 1 Wiring Issue

The overwhelming success of ejection testing indicates that the recovery electronics are

reliable and that the ejection charges are suitable sized for each body tube. Altogether, the

system can be relied upon for triggering recovery events at the appropriate times.

Parachute Deployment Force Testing

The force experienced by the launch vehicle during recovery events was determined by

analyzing acceleration data from the Altus TeleMega. Computing these forces was important for

understanding how the fragile material payload would respond under such conditions and

provided assurance that critical mounting hardware was not in danger of failure.

Using the measured mass of the nosecone/payload section of the launch vehicle and

accelerometer data from the full-scale test flights, it was possible to determine the force exerted

on this section by the recovery harness during various stages of flight. For each flight, the

maximum force occurred during one of the recovery events. These maximum values are given in

Table 54.

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Table 54 - Maximum force on launch vehicle during descent

Flight # Maximum

Force (lbf)

1 198.8

2 95.0

3 206

The forces recorded above are well below the minimum breaking strength of the tubular

nylon recovery harness (4000 lbf) and the tested minimum breaking strength of the recovery

mounting points (insert value here). These results indicate that the launch vehicle is well-

equipped to handle the forces associated with parachute deployment.

Wind Tunnel Testing

Introduction

The wind tunnel is an important instrument used for studying the airflow across solid

specimens. Using a scale model of the rocket inside the wind tunnel for testing helps simulate

the effects of air resistance, or drag force, during the actual flight. The drag coefficient must be

determined in order to best predict the shape, the performance, and the altitude of the rocket. The

experimental drag coefficient will be used to empirically validate simulated CFD and

OpenRocket drag coefficient values.

Testing Apparatus Components

Table 55 shows the apparatus used for performing the wind tunnel experiment.

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Table 55 - Testing Apparatus Components

Instrument

Make/Model

Model number Diameter (in) Length (in)

Width

(in)

Strain gage

Vishay -TN-505-4 -

Strain Gage

- - -

Strain

indicator

Vishay 3800 Wide

Range Strain Indicator

- 0.5 -

Scale Model 0.5

Air fan 144924 10

Wind Tunnel

- Test section

- Motor

-

-

135

-

10 15 -

1295L108A - - -

Differential

pressure

transducer

Honeywell -

SSCSNBN010NDAA5 - - -

Cantilever

beam

6061 rectangle

Aluminum beam

- 7.125 1

Conditioner m-prep conditioner A - - -

Neutralizer

m-prep neutralizer

5A

- - -

139 | P a g e

Instrument

Make/Model

Model number Diameter (in) Length (in)

Width

(in)

Carbide paper

320 and 400 grit

silicon carbide paper

- - -

Degreaser m-prep CSM - - -

Figure 42 through Figure 46 are the components used to set up the experiment. Inside the

wind tunnel, there is an attached electric fan that functions to flow air through the testing area.

When the air crosses the test section, the air pressure increases due to the decrease in cross

sectional area. A pitot tube connected to a differential pressure transducer will be used to

measure the velocity of the air inside the wind tunnel.

Figure 42 – Variable Frequency Drive

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Figure 43- Strain Gage (From Vishay website)

Figure 44 - Strain Indicator

Figure 45 - Air Fan

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Figure 46 - Wind Tunnel

Figure 47 - Example of wiring strain gage to strain indicator

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Figure 48 - Wiring Diagram (strain gage to strain indicator)

The strain indicator will be positioned near the testing area wired with the mounted strain

gage on the 6061 aluminum rectangular beam (refer to Figure 48). There will be a hole in the

test section allowing the operator to insert the beam. The strain gage will be mounted at the base

of the clamped beam. When the air crosses through the test section, the scale model rocket will

resist drag causing deflection in the beam. When the beam is deflected, the strain indicator will

display the strain readings.

- To see how the strain gage wired to the strain indicator, refer to Figure 47 and

Figure 48.

Procedure

1. Strain gage installation.

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1.1.Surface preparation for 6061 Aluminum rectangular beam.

1.2.Degreasing

1.3.Abrading

1.4.Burnishing.

1.5.Conditioning

1.6.Neutralizing.

1.7.Gage bounding

1.8.Apply catalyst

1.9.Apply adhesive

1.10. Soldering strain gage.

1.11. Prepare the leadwire.

1.12. Tin the copper CSA strain gage tabs.

1.13. Trim the lead

1.14. Position the lead wire for soldering.

1.15. Solder the lead wire to the tabs

1.16. Remove all flux residue

1.17. Apply protective coating.

2. Wire the strain gage to the strain indicator.

2.1.Refer Figure 48 to see how strain gage is wired to strain indicator.

2.2.Turn on the strain indicator.

2.3.Set the excitation voltage to be 5 volts

3. Set up the wind tunnel:

3.1.Push the button

144 | P a g e

3.2. Operate the tunnel at airspeeds of 20 mph( 351 in/s) .

3.3.Use the differential pressure transducer to measure the velocity.

4. Use lab view to get the voltage readings.

5. Use equation (4), Equation (5) and Equation (6) to indicate the velocity.

6. Set up the 6061 aluminum rectangular beam:

6.1.Clamp the A 6061 rectangular aluminum beam to the support

6.2.Insert the beam through the test section.

7. Make sure the strain gage is wired to the strain indicator.

8. Position the model rocket inside the test section.

9. Record the reading on the strain indicator readings.

10. Turn off the wind tunnel.

11. Disconnect the strain gage.

Analytical method

From the wind tunnel testing, measured quantities such as velocity and strain will be used

to calculate the drag coefficient. There are two assumptions made before calculating the expected

drag coefficient.

1. The velocity is constant.

2. Air density is constant.

Equation (1) defines the aerodynamic drag coefficient of an object due to air resistance.

CD = Fd

ρ U2

2Ac

(1)

145 | P a g e

Where Fd is the drag force(lbf), ρ is the air density ( lbm/ in3), U is the velocity of the air

wind (in/s), and Ac is the cross sectional area of the scale model rocket. Equation (2) shows the

relation between the strain and the cantilever beam that determines the drag force.

Fd =ϵEwt2

6L (2)

Where ϵ is the strain (in/in), w is the width of the beam (in), t is the thickness of the beam

(in), E is the modulus of elasticity of the beam (lbf/in2), and L is the length where the bounded

gage is positioned (in). Equation (2) is only valid for a rectangular beam. By substituting

equation (2) into equation (1):

CD =Eϵwt2

6Lρ U2

2Ac

(3)

Another measured quantity is the velocity of air. The velocity will be calculated using

the differential pressure transducer. The differential pressure outputs only voltage. Therefore,

there will be at least two related equations for indicating the velocity.

∆P = (Vout − 2.5)5 (4)

Where ∆P is the difference in pressure (in H2O), Vout is the output voltage (volts). In

order to solve for velocity, a unit conversion of the pressure is required.

∆P(

lbf

ft2)= 5.202 ∆P(in H2O) (5)

Therefore, the velocity is calculated using Equation (6).

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U = √2 ∆P(lbf ft2)⁄

0.00226 (6)

Uncertainty

Table 56 shows the mean, systematic uncertainty and random uncertainty used to predict

the total uncertainty of the drag coefficient at a velocity of 352 (in/s). Since the testing is not

performed yet, the strain was calculated using the predicted drag coefficient from the CFD and

OpenRocket simulation. The uncertainty analysis is done for the best case where no precision

error is involved. For the best case uncertainty, the expected total uncertainty for the drag

coefficient was expected to be ±0.097. The Pareto chart (Figure 49) shows the factor that

contributed most to the uncertainty analysis, which is the velocity. Detailed calculations are

provided in Appendix H.

Table 56 - Inputs for Uncertainty analysis

Symbol Description Units Mean Systematic

(Bias)

Uncertainty

Random (Precision)

Uncertainty

L Length in 9.3 0.03 0

b Width in 1 0.0005 0

E Young's Modulus (6061 ALM) psi 10000000 100000 0

t Thickness in 0.1 0.0005 0

u Velocity in/s 352 10.925 0

∈ Strain in/in 0.000171569 2.06E-06 0

ρ Density lb/in3 0.0004 0 0

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Ac Area of the subscale rocket. in2 2.405281875 0.0005 0

Figure 49 - Pareto Chart

Test Status

The wind tunnel test has not been performed yet. A thorough uncertainty analysis was

performed before proceeding with the test procedure to address concerns of accuracy.

Complicating factors included:

Test section size (length and diameter of the scale model must be reduced due to minimal

size constraints of the test section in the wind tunnel)

Surface finish of the 3D model (surface roughness on minimal scale could affect

coefficient of drag)

Cantilever beam assumption validity

0.00

10.00

20.00

30.00

40.00

50.00

60.00

70.00

80.00

90.00

100.00

Pe

rce

nt

(%)

Systematic (Bias)(%)

Random (Precision)(%)

148 | P a g e

Uncertainty of equipment

Validity of results as check for CFD

Scale Model Testing

The sub-scale model was tested in December. It had a goal to reach an apogee of 2,500 feet.

The model was launch twice successfully. Although the first flight reached an apogee of 2592

feet, we learned that we were not using the correct black powder or enough black powder.

Changing the black powder for the second flight resulted in no issues on the second sub-scale

flight. The second flight was closer to our target and reached an apogee of 2498 feet. For a more

detailed breakdown, refer to the CDR report.

Payload Testing

Before the entire payload assembly was tested, the spring constant for the 5 base springs

given by the manufacturer was tested to verify that the values used in the math model were

accurate, (for math model refer to CDR).. No variable can affect the tests except the change in

weights. After these criteria where met, the test was deemed successful if the spring deforms,

measurements are accurately taken, and the weight is properly recorded. The procedure for the

spring constant test is as follows.

1. Fasten the spring to a mounting plate and turn apparatus upside down so that weights can

be suspended from it

2. Fashion a hook and attach it to the end of the spring so that weights can be attached

3. Measure and record the un-stretched length of the spring

4. Attach a weight to the hook

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5. Record the mass of the weight and change in length of the spring

6. Repeat steps 4 and 5 until enough data has been collected

7. Calculate the spring constant using Hooke’s Law, Equation 1.

𝑭 = −𝒌𝒙 Equation 1: Hooke’s Law

The results of this test are summarized in Table 57.

Table 57 - Spring Constant Test Values

Mass

(kg)

Weight

(lbf)

Spring

Displacement

(m)

Spring

Displacement

(in)

k (kg/m) k (lbf/in)

2 4.4 0.00635 0.25 314.961 17.6

4 8.8 0.015875 0.625 251.969 14.08

6 13.2 0.0254 1 236.220 13.2

8 17.6 0.03175 1.25 251.969 14.08

10 22.0 0.034925 1.375 286.328 16

12 26.5 0.041275 1.625 290.723 16.308

15 33.1 0.053975 2.125 277.906 15.576

Average 272.869 15.263

Uncertainty ±.16 (kg/m) ±.23 (lbf/in)

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The manufacturer specified spring constant is 15 confirming that the tested springs were

reasonably close to the specified values. The weights picked for the spring constant test started

at 2 pounds and went incrementally until the spring failed to get the entire range of forces.

After confirming that the spring constant was near what the manufacturer had specified,

a drop test was performed on the payload assembly while mounted in a mock rocket body tube

created out of Blue tube. The test was a vertical drop from three stories or 30 feet high. This

was designed to simulate forces worse that actual flight and to calculate the force endured by the

fragile material within Cylinder 1 through the use of an accelerometer. After the first test

however, three of the base springs epoxy and welds broke causing the springs to buckle and five

of the wire rope isolators failed, three had adhesive failure and two had cohesive failure. One

reason for the failure was that the math model simulated the force as a purely longitudinal force

along the length of the rocket, however during the drop test, the tube hit the ground at

approximately a 45 degree angle. The first drop test was meant to determine the amount of

freefall time to properly calculate the impact force and acceleration that the payload

experienced. However, due to the failure of the springs and a malfunction with the

accelerometer, no data was gathered. Due to the drop test’s lack of repeatability, a substitute test

was designed with maximum variable control providing more accurate data. This test was the

modified Charpy Impact test.

The modified Charpy impact test employed for lack of sources of error and repeatability.

The test was set up by placing the payload assembly in the mock rocket body tube and

positioning it in the impact zone of the hammer on the Charpy Impact test machine. The U-bolt

used to attach the parachute was also used to be the connection point where the hammer

transmitted its force to the payload. A frame with sheets draped over it was set at the end of the

151 | P a g e

testing apparatus to catch the payload when it was launched from the machine. Several tests

were conducted to determine the reduction in acceleration, the optimal fill material, and overall

performance of the payload. The testing data can be seen in Table 58. To be able to calculate the

percent reduction in force and acceleration, the accelerometer was first mounted to the outside of

Cylinder 2 or on the mock rocket body to get a base acceleration value. This was later used in

comparison to the accelerometer values within Cylinder 1 showing the percent reduction.

Table 58 - Charpy Impact Acceleration Test Data

Initially, the fill material selection was going to be based off the accelerometer values and

percent reduction given from those. However, as can be seen in Table 58, the acceleration values

were very different between the tests using shredded paper and cotton filling. The values were

determined by the graphs found in Appendix I – Payload Accelerometer Graphs. By the time this

had been discovered, the testing housing had been disassembled and the payload was already in

use in the full-scale rocket, so further testing could not be done to determine the reason for such

a large difference in acceleration. The “Base Value” seen in the table represents the acceleration

values in the x, y and z directions for the accelerometer mounted to the mock body tube

receiving acceleration reducing effects of the springs or fill material. This was used as the basis

for comparison. For the shredded paper and cotton fill tests, the accelerometer was placed inside

Cylinder 1 to mimic what the fragile material will endure. One source of error and possible

Fill Material

Cotton Filling Shredded Paper Paper/Cotton mix Base Value

Acceleration

(counts/g) average

of x,y and z

directions.

11444 2452 N/A 11643

Percent reduction

from base1.7 78.9 N/A N/A

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explanation of why the accelerometer values were so different could be the rate of data logging.

The maximum rate of data logging for the model of accelerometer used was 4 hz. This means if

the impact occurred in a small enough time step, the entire event could have been missed and not

logged, which is most likely what happened in all 3 tests. The systematic uncertainty for the

accelerometer used is given as:

Nonlinearity (x,y,z)=±0.5% 𝐹𝑆 ; Where FS= 32 g Equation 2

Nonlinearity (x,y,z)=±0.005 × 32 = ±0.16𝑔 Equation 3

Zero-g Offset level accuracy:

X and Y-Axis = ±150 mg= ±0.15 g. Equation 4

Z- Axis = ±250 mg= ±0.25 g. Equation 5

Overall systematic error = B = BOIE=√𝑒𝐿2 + 𝑒𝑧

2 Equation 6

BOIE(x,y) =√𝑒𝐿2 + 𝑒𝑧

2= √(0.16𝑔)2 + (0.15𝑔)2 =±0.2193 𝑔 Equation 7

BOIE(z) =√𝑒𝐿2 + 𝑒𝑧

2= √(0.16𝑔)2 + (0.25𝑔)2 =±0.2968 𝑔 Equation 8

Although the systematic uncertainty demonstrates accuracy in the accelerometer measurements

the decision was made to base fill material choice off of the survival of the sample fragile

material specimens shown in Table 58. However, this means that no numerical data can prove

that the payload reduces the maximum force and acceleration felt by the fragile material by at

least 50 percent.

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The fragile material tested in the Modified Charpy Impact test was tested first with the

frame and draped sheets to catch the payload after impact. The testing data can be seen in Table

59.

Table 59 - Fragile Material Sample Testing

The original test matrix in the CDR included several other fill materials to test, however other

materials were omitted due to volume density considerations. Each fragile material was first

tested with the hammer of the Charpy Impact Tester at 90 degrees or parallel with the floor, and

all fragile materials survived in each of the fill materials. After no fragile material had broken,

each fragile material was then tested with no fill material. From the testing with no fill material,

the egg was determined to be the most fragile of all materials. The same egg was re-tested 2

times with the hammer on the Charpy Impact Tester raised to the maximum as in every test.

However, this time a plywood board supported with cinder blocks was placed 2 feet from the

payload so that immediately after impact with the hammer, the payload would impact with a

Break y/n? Fill Material

Cotton Filling Shredded Paper Paper/Cotton mix no fill

2 large incadescent bulbs no no no no

2 candelobra bulbs no no no yes

Fragile material glass sheet no no no yes

egg 1/2 power swing no no no yes

egg full power swing no no no yes

egg full power swing double

impact no no no yes

154 | P a g e

sturdy wall. The wall allowed a simulation of compression as well as tension in all springs for

one test. This test was performed only with the egg as the fragile material selection and both

times the egg survived un-cracked.

Since both the shredded paper and cotton fill worked in protecting the fragile objects, a

combination was selected for the final payload design. Shredded paper will be placed in the top

and bottom of the cylinder to crumple and provide axial cushion while the fragile object will be

wrapped in cotton fill to project a majority of side impact and keep the material centrally located

in Cylinder 1.

The final test the payload endured was the full-scale flight tests. The rocket was flown

three times, and each time an egg was placed in the payload with the shredded paper and cotton

fill mix. During the first two flights, the accelerometer was placed in Cylinder 1 with the egg to

try to obtain the maximum force experienced by the fragile material. Accelerometer data for

Flight 1 is seen in Figure 50.

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Figure 50 - Accelerometer Data Full-scale Flight 1

Flight 1 was the only full scale flight that the accelerometer recorded data for due to the

battery malfunctioning. The graph shows the maximum acceleration in the x direction which is

the direction of flight of the rocket. This value is 32307 counts/Hz and converting to acceleration

values is 507.95 ft/sec2. The altimeter data from the scoring altimeter shows the maximum

acceleration of the entire rocket as being 443.5 ft/sec2. The uncertainty for the both acceleration

values are under ±10ft/sec2 the proving that the accelerometer used malfunctioned during

recording. This also helps prove the random values that occurred during the Charpy Impact

Tests. Also through all 3 test flights, the same egg was used and each flight the egg survived

unscathed. The team considered this to be evidence of successful performance of the fragile

material payload.

-40000

-30000

-20000

-10000

0

10000

20000

30000

40000

0 200 400 600 800 1000 1200 1400 1600

Acc

eler

atio

n (

cou

nts

/hz)

Time Step 4 Hz

X Y Z

156 | P a g e

Full Scale Testing

The full-scale rocket was tested on February eighteenth. A successful flight was defined by,

the fragile material payload needed to survive the entire flight, and the apogee needed to be

within 5,125 feet and 5,375 feet. Table 60 shows a summary of the results from the full scale

flights. A full review of the full scale test can be found in the Full Scale Flight Analysis section

above.

Table 60 - Full Scale Flight Results

Apogee Did the Payload Survive?

Flight 1 4913 feet Yes

Flight 2 4795 feet Yes

Flight 3 5291 feet Yes

Requirements Compliance

In order to be succeed in the competition, and follow all rules and regulations set forth by

NASA, the team will abide by both NASA & team-created requirements. These requirements

involve various facets of the project from rocket design parameters, to launching procedures, and

safety protocols. Each individual NASA requirement is listed in Table 61, sorted by the

corresponding USLI Handbook number. Within this table, each requirement is summarized and a

verification plan is given to ensure compliance with all NASA requirements. Additionally,

information has been added pertaining to the status of each item, as well as where further

information can be found in this report.

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Table 61 - NASA Requirement Compliance

NASA Requirements

Handbook

Number Summarized Requirement

Verification

Method(s) Description of Verification Plan

Status & Location

1.1

The vehicle shall deliver the

science or engineering payload

to an apogee altitude of 5,280

feet above ground level (AGL).

Test

Analysis

The rocket team will utilize

OpenRocket, RockSim, CFD, &

test flight data to achieve an

accurate prediction of altitude.

Full scale test completed with

apogee of 5,291 feet. See “Full

Scale Flight” for more detail.

1.2

The vehicle shall carry one

commercially available,

barometric altimeter for

recording the official altitude

used in determining the altitude

award winner.

Inspection

The rocket will house a Atlus

Metrum TeleMega altimeter in

the nosecone to record the official

altitude used in determining the

altitude award winner.

Three altimeters meeting

requirements were flown for

full scale; all producing valid

altitudes. See “Validity

Assessment” for more detail.

1.3

All recovery electronics shall be

powered by commercially

available batteries.

Inspection Batteries & altimeter will be

purchased from online rocketry

sources.

Recovery altimeters powered by

Energizer 9V Lithium Batteries.

See “Line Item Budget in

Appendix F”.

1.4

The launch vehicle shall be

designed to be recoverable and

reusable. Reusable is defined as

being able to launch again on the

same day without repairs or

modifications.

Test

Inspection

The rocket is reusable in design

because the team is using a motor

that has refuels that can be

reloaded into the motor under

supervision.

Three test flights were

conducted on the February 18th.

No repairs were made to the

rocket, making it reusable. See

“Full Scale Flight”.

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NASA Requirements

Handbook

Number Summarized Requirement

Verification

Method(s) Description of Verification Plan

Status & Location

1.5

The launch vehicle shall have a

maximum of four (4)

independent sections.

Inspection

The launch vehicle will have 3

independent sections: the aft body

tube, the bow body tube and

nosecone, and the coupler.

The launch vehicle has 3

independent sections: the aft

body tube, the bow body tube

and nosecone, and the coupler.

See “Vehicle Criteria”.

1.6 The launch vehicle shall be

limited to a single stage.

Inspection

Demonstration

The launch vehicle shall be a

single stage.

Only one L850W is used, as

seen in “Vehicle Criteria”.

1.7

The launch vehicle shall be

capable of being prepared for

flight at the launch site within 4

hours.

Test The team will conduct multiple

tests on full-scale test day and

measure re-launch times.

The team was able to prepare

the rocket in 32 minutes on

February 18th. See “Full Scale

Flight” for more detail on

multiple launches that day.

1.8

The launch vehicle shall be

capable of remaining in launch-

ready configuration at the pad for

a minimum of 1 hour without

losing the functionality of any

critical on-board component.

Test

The launch vehicle design will

ensure all components have a life

of greater than 1 hour without

loss of functionality via a full-

scale launch pad test.

All systems remained on during

the full scale test for over 2

hours while the rocket was

stuck in a tree. See “Full Scale

Flight” for more detail.

1.9

The launch vehicle shall be

capable of being launched by a

standard 12-volt direct current

firing system.

Inspection

Test

The ignition system will use a 12-

volt direct current firing system.

The ignition system and igniters

used during the full-scale test is

12V. See the “Line Item

Budget” in the Appendix for

exact ignitor specifications.

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NASA Requirements

Handbook

Number Summarized Requirement

Verification

Method(s) Description of Verification Plan

Status & Location

1.10

The launch vehicle shall require

no external circuitry or special

ground support equipment to

initiate launch (other than what is

provided by Range Services).

Inspection

There will be no external circuity

for the ignition system because it

will be a ground based ignition

system being placed underneath

the rocket before launch with 300

ft of cord between the igniter and

the controller.

The ignition system comprised

of only 1 ignitor that runs off of

12V. This setup was successful

during the full scale test. See

the “Line Item Budget” in the

Appendix for exact ignitor

specifications.

1.11

The launch vehicle shall use a

commercially available solid

motor propulsion system using

ammonium perchlorate

composite propellant (APCP)

which is approved and certified

by the National Association of

Rocketry (NAR).

Inspection

The motor being used is a solid

fuel motor from AeroTech. The

motor is the L850W.

The team has purchased and

flown on an Aerotech L850W,

see the “Line Item Budget” for

further information on the

motor.

1.12 Pressure vessels on the vehicle

shall be approved by the RSO. Inspection No pressure vessels will be used.

As of final design, no pressure

vessels are used. See “Design

and Construction of Vehicle”.

1.13

The total impulse provided by a

University launch vehicle shall

not exceed 5,120 Newton-

seconds (L-class).

Inspection

The motor will produce an

impulse of 3695 N-s which is

below the specified total impulse

that is allowed.

The motor details can be found

via Aerotech’s website. See

“Line Item Budget” for specific

motor look-up information.

1.14

The launch vehicle shall have a

minimum static stability margin

of 2.0 at the point of rail exit.

Test

Analysis

Using OpenRocket, Rocksim, and

Test Data – determine rail exit

velocity and then stability.

The flight configuration for

competition has an actual flight

stability of 2.70.

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NASA Requirements

Handbook

Number Summarized Requirement

Verification

Method(s) Description of Verification Plan

Status & Location

1.15

The launch vehicle shall

accelerate to a minimum velocity

of 52 fps at rail exit.

Analysis

The rocket team will utilize

OpenRocket, RockSim, CFD, &

test flight data to achieve an

accurate prediction of minimum

velocity at rail exit. The current

value is 66.9 fps.

The Full-Scale flight was a

success. See Section “Flight

Simulations and Altitude

Predictions”

1.16

All teams shall successfully

launch and recover a sub-scale

model of their rocket prior to

CDR.

Test

A sub-scale model with

comparable weights, lengths, and

masses will be launched prior to

the CDR.

The Sub-Scale test was

successful and has been

completed. See the CDR’s

“Sub-Scale” Flight section.

1.17

All teams shall successfully

launch and recover their full-

scale rocket prior to FRR in its

final flight con- figuration.

Test The project schedule will ensure a

full-scale rocket launch occurs

before the FRR.

Launch both complete and

successful on February 18th.

See “Full Scale Flight” for more

detail.

1.18

Any structural protuberance on

the rocket shall be located aft of

the burnout center of gravity.

Test

Analysis

No structural protuberances will

exist bow of the burnout center of

gravity.

The rocket has 3 bolts holding

the nosecone to the bow body

tube. These are located bow of

the burnout center of gravity but

has been cleared by NASA. No

other structural protuberances

exist bow of the burnout center

of gravity.

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NASA Requirements

Handbook

Number Summarized Requirement

Verification

Method(s) Description of Verification Plan

Status & Location

1.19 Vehicle Prohibitions Inspection

The launch vehicle will follow all

prohibitions laid out in section

1.19 of the 2017 SL NASA

Student Handbook.

The launch vehicle has

followed all prohibitions laid

out in section 1.19 of the 2017

SL NASA Student Handbook.

See “Vehicle Criteria” for full

design.

2.1

Vehicle must deploy a drogue

parachute at apogee, followed by

a main parachute at a much

lower altitude.

Test Dual-deployment altimeters are

programmed to fire ejection

charges at apogee and at 750 feet.

Full-scale test flights resulted in

successful recovery events. See

“Full Scale Flight” &

“Recovery” for more.

2.2

A successful ground ejection test

for both parachutes must be

conducted prior to sub- and full-

scale launches.

Test Multiple ejection tests conducted

prior to sub- and full-scale

launches.

Sub-scale and full-scale test

ejections were successful – 8

consecutive full scale test

ejections. See “Ejection

Testing” section.

2.3

No part of the launch vehicle

may have a kinetic energy

greater than 75 ft-lbf at landing.

Analysis

Demonstration

Parachute sizes are optimized to

minimize kinetic energy at

ground impact.

Full-scale test flights resulted in

kinetic energy below the

maximum allowable. See

“Recovery” section.

2.4

Recovery electrical circuits must

be independent of payload

circuits.

Inspection Recovery electronics are housed

in a separate compartment.

Coupling tube constructed

completely independent of other

electronics. See “Recovery”

section for more detail.

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NASA Requirements

Handbook

Number Summarized Requirement

Verification

Method(s) Description of Verification Plan

Status & Location

2.5

Recovery system must include

redundant, commercial

altimeters.

Inspection Two PerfectFlite Stratologger CF

altimeters will be used.

Redundant ejection charges

observed for each recovery

event during full-scale test

flights. See “Recovery”

section.

2.6

Motor ejection cannot be used

for primary or secondary

deployment.

Demonstration

Inspection

Black powder ejection charges

are used to deploy parachutes.

Black powder ejection charges

successfully triggered recovery

events for full-scale test flights.

See “Recovery” section.

2.7

Each altimeter must be armed by

a dedicated switch accessible

from the rocket exterior.

Inspection A separate switch accessible

through pressure sampling holes

is used to arm each altimeter.

Rotary switches successfully

armed from rocket exterior for

full-scale test flights. See

“Recovery” section.

2.8 Each altimeter must have a

dedicated power supply. Inspection

Separate 9-Volt batteries are

wired to the power leads of each

altimeter.

Recovery altimeters were

powered up for duration of each

full-scale test flights. See

“Recovery” section.

2.9 Each arming switch must be

lockable to the “ON” position. Inspection

Locking rotary switches are wired

to the switch leads of each

altimeter.

Recovery altimeters were

powered up for duration of each

full-scale test flights. See

“Recovery” section.

2.10

Removable shear pins must be

used to seal the parachute

compartments.

Inspection Three #2 nylon shear pins are

used to seal each parachute

compartments.

Pins sheared successfully

during ejection testing and full-

scale test flights. See

“Recovery” section.

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NASA Requirements

Handbook

Number Summarized Requirement

Verification

Method(s) Description of Verification Plan

Status & Location

2.11

Tracking device(s) must transmit

the position of any parts of the

launch vehicle to a ground

receiver.

Test

Demonstration

Inspection

All parts of the launch vehicle are

tethered together; position will be

transmitted via a flight computer

in the nosecone.

All sections of launch vehicle

remained tethered during full-

scale test flights. Position data

was successfully transmitted

throughout each flight. See

“Line Item Budget” for exact

GPS specifications.

2.12

Recovery system electronics

must not be adversely affected

by any other on-board

electronics.

Test

Inspection

Recovery electronics located in

separate compartment.

Recovery altimeter data showed

no signs of interference after

full-scale test flights. See

“Recovery” section.

3.4.1

Design container capable of

protecting an unknown object of

unknown size and shape.

Testing

Math model is used to develop

spring system in conjunction with

a concentric cylinder model to

provide sufficient vibration

dampening and force reduction.

Full scale flights resulted in safe

return of an egg – with ability to

adjust to multiple eggs. See the

“Payload Testing” section for

more detail, including a %

reduction in force.

3.4.1.2 Object must survive duration of

flight Testing

The spring and concentric

cylinder design will be tested

with a matrix of different support

materials as well as testing

materials to assure the unknown

object(s) can survive the flight

during demonstration.

Full scale flights resulted in safe

return of an egg. See the

“Payload Testing” section for

more detail, including a %

reduction in force.

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NASA Requirements

Handbook

Number Summarized Requirement

Verification

Method(s) Description of Verification Plan

Status & Location

3.4.1.4

Once the object is obtained, it

must be sealed in its housing

until after the launch and no

excess material may be added

after receiving the object.

Test During full scale flight, verify

that an object can be contained

using no excess material.

Using only material already in

the rocket, this setup was tested

on February 18th for the full

scale flight and passed. See

“Payload Testing” section for

more.

4.1 Each team shall use a launch and

safety checklist

Inspection

Demonstration

Final assembly and pre-launch

checklists will be created and

reviewed at the appropriate time

to ensure safe launch of the rocket

and all members involved in the

launch

Launch and safety checklist

used for full-scale test flight.

See “Launch Operations”

section for more detail.

4.2

Each team shall identify a

student safety officer who shall

be responsible for the safety of

the team and ensure all proper

rules and guidelines are followed

Inspection

The team has appointed a safety

officer to monitor the safety of

the team throughout the project

and ensure all federal rules and

laws are met.

Safety officer Bryan Bauer

oversaw both fabrication and

testing phases to ensure safe

and successful operations.

4.3

The team safety officer shall

monitor team activities with an

emphasis on safety throughout

the design, construction, and

testing of the rocket by

maintaining MSDS sheets and

hazard analyses.

Inspection

The team safety officer will

monitor the progress of the

project emphasizing the proper

safety procedures for the current

stage of the project.

Safety officer has monitored the

full-scale testing, fabrication

and launch in order to ensure

safe operations.

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NASA Requirements

Handbook

Number Summarized Requirement

Verification

Method(s) Description of Verification Plan

Status & Location

4.4

Each team shall appoint a mentor

who has certification and is in

good standing with the NRA.

Inspection

The team has assigned an school

faculty member to mentor the

project to provide valuable insight

on the rocket design and

construction as well as assume

full liability of the rocket.

Dr. David Unger is the team

mentor; his information can be

found on the cover page.

4.5

During test flights, teams shall

abide by the rules and guidance

of the local rocketry club's RSO

Demonstration

Team will converse with RSO at

local rocketry club to ensure all of

their chapter’s rules and

regulations are abided by.

Team is in compliance will all

rules and regulations set forth

by local rocketry club

“BluesRocks”. See “Full Scale

Flight” for more detail.

4.6 Teams shall abide by all rules set

forth by the FAA Demonstration

Team will converse with NASA

lead safety officer and thoroughly

research all rules and regulations

set forth by the FAA to ensure all

rules and regulations are abided

by.

Team is in compliance will all

rules and regulations set forth

by FAA and NASA. See “Full

Scale Flight” for more detail on

the flight.

5.1

Students shall do 100% of the

project excluding motor / black

powder handling.

Demonstration

Inspection

The team will continuously

demonstrate an independently

managed and executed project.

The team lead will routinely

monitor this quality.

The team has only used mentors

for guidance and will continue

to do so.

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NASA Requirements

Handbook

Number Summarized Requirement

Verification

Method(s) Description of Verification Plan

Status & Location

5.2 A detailed project plan shall be

maintained. Demonstration

Documents for scheduling,

budget tracking, outreach, and

safety will be continuously

updated and reported.

Project plan is updated and can

be found in the “Project Plan”

section.

5.3 Foreign National members shall

be identified by the PDR. Inspection

The team lead will ensure that

any Foreign National members

are clearly indicated in the PDR.

Foreign National members have

been identified in emails with

NASA.

5.4

All team members attending

launch week shall be identified

by the CDR.

Inspection

It will be checked that a list of

team members, with indications

of those attending launch week,

will be included in the CDR.

Team members have been

identified in emails with NASA,

along with completed waivers.

5.5

The educational engagement

requirement shall be met by the

FRR.

Inspection

The Educational Engagement

lead shall confirm that all

documentation has been received

and approved by NASA prior to

the FRR.

Team has completed outreach

activities with over 200 students

reached. Educational

engagement is not discussed in

this report.

5.6 The team shall develop and host

a website for documentation. Test

Team members will periodically

confirm that the website is

functioning as intended by

opening each posted document.

Website has been developed

and is being updated.

5.7

The team shall post & make

available for download all

deliverables by the specified

date.

Inspection The team lead shall confirm that

all documents are posted prior to

the specified date.

All reports, presentation slides

and flysheets have been and

will continue to be posted to the

team website by the deadline set

forth by NASA.

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NASA Requirements

Handbook

Number Summarized Requirement

Verification

Method(s) Description of Verification Plan

Status & Location

5.8 All deliverables must be in PDF

format. Inspection

The team lead shall confirm that

all documents posted are in PDF

format.

All deliverables to the team

website are upload in PDF

format.

5.9 A table of contents must be

included in all reports. Inspection

The team lead shall ensure that a

table of contents is located at the

start of each report.

See Table of Contents, Figures,

and Tables sections.

5.10 Page numbers shall be provided

in each report. Test

Page numbers shall be checked to

the table of contents to ensure

continuity throughout the report.

See lower right hand corner of

each report.

5.11

The team shall provide

videoconference equipment

needed for reviews.

Demonstration

Test

Videoconference rooms will be

reserved and trialed immediately

prior to each design review.

Requirement met, same setup

will be used for all future

correspondence.

5.12 All teams shall use launch pads

provided by the SLS provider. Demonstration

The team shall design the rocket

to utilize 1515 12’ launch rail.

The 12’ 1515 Rail used for sub-

scale launch operated as

intended, see Full Scale Flight

section.

5.13 The team must implement the

EIT accessibility standards. Demonstration

If software or applications are

created (not planned) the team

will abide by 36 CFR Part 1194.

Otherwise, all components

containing software will be

checked to ensure compliance.

Software not designed by team.

See “Line Item Budget” for

exact electronic components.

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As mentioned, Project ACE has developed a set of team derived requirements as well. The

team requirements can be seen in Table 62. They cover things that were not touched on by the

handbook and also add depth to certain handbook requirements.

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Table 62 - Team Requirement Compliance

Team Requirements

Number Requirement Verification

Method

Description of Verification

Method

Status & Location

1

All reports shall be compiled at

least three days prior to NASA

due dates.

Demonstration

Reports shall be completed,

according to team schedule, prior

to NASA due dates to allow for

revision time and mitigate risk of

late submissions.

The team has completed all

reports on time. The dates can

be seen in the “Schedule”

portion of the report.

2

Each member of the team shall

have a working knowledge of

each subsystem.

Inspection

At each team meeting, every sub-

section lead will review the status

of their section with the entire

team. The team leader will

confirm that the information

presented is sufficient.

This has been maintained. It

was recently demonstrated at

the full-scale launch where

team members had to work on

each other’s sections. See “Full

Scale Flight” for more details.

3 Safety shall be made the team’s

first priority. Demonstration

The safety officer will

periodically ask team members

what the most important aspect of

the project is.

Safety officer has asked 17

team members what the most

important part of the project is

and has had 15 “safety”

answers. The two outliers have

been reminded of safety.

4 Altimeters shall be in good

working order. Test

All altimeters shall be flown on

sub-scale and full scale flight

tests. Altitude readings will be

compared to confirm consistency.

Altimeters have all been

extensively tested and have

passed all tests. See “Altimeter

Testing” section for more detail.

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Team Requirements

Number Requirement Verification

Method

Description of Verification

Method

Status & Location

5 The tracking system shall be in

good working order. Test

The tracking system shall be

flown on the sub-scale and full

scale flight tests. This will be

used to find the rockets thus

confirming its operation.

The tracking system performed

perfectly in the sub-scale &

full-scale test. During the full

scale test, the tracking system

located the rocket over 1 mile

away.

6 A solid output signal must be

given from triggered altimeters.

Test

Analysis

All altimeters will be triggered

while voltage is read on the

output. This output will be read

to confirm it is acceptable.

The output voltage is seen in

real time at the base station.

7 All circuits shall be checked

prior to use. Inspection

All circuits will be confirmed at

each node to ensure connections.

This was completed for both

sub-scale and full-scale tests.

Continuity and amperage were

both inspected.

8

Impulse for the parachute

deployment shall be determined

experimentally.

Test

Analysis

The main parachute shall have an

apparatus (strain gauge) attached

to it that enables a force to be

read as it opens at high speed.

This will cut down in the large

ambiguity that exists in

estimating an impulse value.

Parachute force testing

completed using acceleration

data on altimeter. See

“Parachute Deployment Force

Testing” section.

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Team Requirements

Number Requirement Verification

Method

Description of Verification

Method

Status & Location

9

A spring constant for parachute

cords shall be determined

experimentally.

Test

Analysis

The spring constant shall be

determined using forces related to

what is experienced with

parachute opening. This helps

when estimating energy

absorption by the cord when the

chute opens.

This testing has been bundled

into the parachute deployment

force test. The parachute as it

relates to the body had its

acceleration measured, which as

a system includes the cords’

expansion. See “Parachute

Deployment Force Testing”.

10 Payload must reduce force felt

by object(s) by 50 % Testing

From the mathematical model,

appropriate springs will be

selected to induce oscillation and

reduce force. These will be tested

by Charpy Impact Tests.

This testing has been completed

and selected springs were also

tested to assure spring constants

given by the manufacturer were

accurate.

11 Payload must reduce

acceleration of object(s) by 35 % Testing

From the mathematical model,

appropriate springs will be

selected to insure acceleration

graphs show 35 percent reduction

from inputs. Will be tested via

Charpy Impact Test.

The test to deduce the percent

reduction in force and

acceleration was completed

however a faulty accelerometer

caused data to be useless and

therefore percent reduction

cannot be found. See Payload

Testing section for full

explanation.

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Team Requirements

Number Requirement Verification

Method

Description of Verification

Method

Status & Location

12 Electronics must operate in cold

temperatures

Demonstration

Testing

First, temperature sensitive

components will be identified.

Then components will be tested

in the cold with ejection testing.

2 Full Scale ejection tests done

in cold (-2°C) environment, test

passed. See “Ejection Testing”

section.

13 Mach Number will be less than

0.6

Simulation

Test

From simulations, the motor and

aerodynamics of the rocket will

ensure the rocket has a Mach

number of 0.53

Full-Scale completed and

simulations ran. See section

“Flight Simulations and

Altitude Predictions”

14 The rail exit velocity will be

above 60 ft/s Simulation

From the simulations, the rocket

weight and motor section will

ensure of having a rail exit

velocity of 66.9 ft/s

Full-Scale completed and

simulations ran. See section

“Flight Simulations and

Altitude Predictions”

15

Complete a Combustion

Analysis on the Motor to obtain

Pressure of fuel ignition

Simulation

From hand calculations to obtain

the temperature and pressures to

run the FEA analysis on the

motor casing to find the Factor of

Safety of the motor casing

Combustion analysis complete

and has a pressure of 2155 kPa.

Located in Combustion

Analysis in CDR.

16

Complete a Modal Analysis on

the Motor Mount System to

ensure safety and stability of the

rocket

Simulation

Used hand calculations to

determine the natural frequency

of the motor mount and then used

Finite Element Analysis to find

operational frequency

Modal analysis complete with

an operating frequency not near

natural frequency. See “Modal

Analysis” in CDR.

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Team Requirements

Number Requirement Verification

Method

Description of Verification

Method

Status & Location

17

Complete a Shear Stress

Analysis on the motor mount to

ensure that the epoxy being used

will withstand the motor forces

Demonstration

Used hand calculations to have a

verification of the results found

using the Finite Element Analysis

to find the Factor of Safety of the

motor mount

Analysis complete with F.O.S.

of 17.07. See “Shear Stress

Analysis” in CDR.

18

Have a Factor of Safety above 2

for the Combustion Analysis and

Shear Stress Analysis

Demonstration

Calculated the Factor of Safety of

Combustion and Shear Stress

areas. Found Combustion factor

of Safety to be 103. Found Shear

Stress factor of safety to be 17

Complete with F.O.S. of 103.6

& 17.07, respectively. See

“Propulsion” Section in CDR.

19 Reach an altitude between 5,200

and 5,400 feet

Simulation

Testing

Use OpenRocket and Rocksim to

simulate the altitude of the full-

scale rocket. Test the full-scale to

see the actual altitude

Full scale test apogee of 5,291

feet. See “Full Scale Flight”

section for more detail.

20

Design flexibility on full-scale

test launch day to raise or lower

altitude on a second test-flight.

Demonstration

Simulation

Using simulation and

demonstration of design, the team

will prove that on test launch day,

small changes can be made to

raise or lower altitude for a

second flight.

Ballast adjusted on February

18th to change height for three

different flights. See “Full

Scale Flight” for more detail on

the configurations.

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Budgeting and Timeline

Budget

Project ACE received funding from three primary sources. First, the Indiana Space Grant

Consortium generously awarded Dr. David Unger & Project ACE a total of $5,000.00. The

University of Evansville’s Student Government Association (SGA) and University of

Evansville’s College of Engineering & Computer Science contributed as well, resulting in total

funding of $10,530.00 - seen in Table 63.

Table 63 - Sources of Funding

Source Amount

NASA Grant (INSGC) $5,000.00

Student Government Association $2,730.00

U.E. College of Engineering & Computer Science $2,800.00

Total $10,530.00

After obtaining funding, Project ACE created a detailed budget that resulted from a

complete parts list (Appendix F). For financial purposes, this budget broke the project into 10

sections. Additionally, a variable contingency fund was built into the budget for each section.

The sum of the parts list and variable contingency fund is shown as the “Forecasted Amount”

column in Table 64.

Using detailed cost-tracking methods an “Amount Expended” column was created in

Table 64. The Amount Expended figure represents the total amount spent on that section of the

project. As of FRR submission, all spending has been completed aside from fuel to/from

competition. Fuel costs have been conservatively estimated and are included in the “Travel /

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Lodging” figures. As such, all expended amounts reflect final values. From this a “Difference”

column was created – that is the difference between the forecasted and expended amounts.

Figures containing parenthesis and a red background indicate a section that went over budget

while figures with a green background indicate a section that remained under budget. A visual

comparison of forecasted and actual expenses is provided in Figure 51.

Table 64 - Sectional Budget Breakdown

Operating costs were over budget due to the purchase of tools and team polo

reimbursements. The main payload went over budget due to the mounting re-design (discussed

in Payload section) & while the recovery excess was caused by unforeseen component costs.

Section Forecasted Amount Amount Expended Difference

Operating $300.00 $570.90 $(270.90)

Travel / Lodging $2,730.00 $2,475.51 $254.49

Launch Pad $220.00 $197.59 $22.41

Aerodynamics (Body) $1,400.00 $962.98 $437.02

Propulsion $2,500.00 $2,235.14 $264.86

Main Payload $500.00 $792.17 $(292.17)

Electronic Payload $630.00 $614.86 $15.14

Recovery $1,150.00 $1,189.49 $(39.49)

Scale Model $1,000.00 $993.05 $6.95

Educational Engagement $100.00 $74.83 $25.17

Total $10,530.00 $10,106.52 $423.48

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Ultimately, the project concluded under budget by $423.48 – a total expenditure nearly 5% under

the forecast.

Figure 51 - Sectional Budget Amounts

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Schedule

A detailed breakdown of each task, accompanied with all pertinent dates, can be found in the

detailed task breakdown in Appendix G. All critical dates for completion of the project are

shown in Table 65. Additionally, a broader view of the task breakdown can be seen in Gantt

chart form in Figure 52. Despite a few testing delays the project is on schedule as of FRR.

Table 65 - Critical Dates

Activity Due Date

NASA U.E. Team

Project Kickoff Aug. 15, 2016 - -

General Motor Selection/Data Sept. 30, 2016 - Sept. 16, 2016

Informal Design Sketches - Sept. 21, 2016 Sept. 19, 2016

Proposal Sept. 30, 2016 Oct. 3, 2016 Sept. 27, 2016

Motor Selection/ Data Oct. 31, 2016 Oct. 7, 2016

Proposal Presentation - Oct. 24, 2016 Oct. 22, 2016

PDR Report Nov. 04, 2016 - Oct. 26, 2016

PDR Flysheet Nov. 04, 2016 - Oct. 26, 2016

PDR Presentation Nov. 04, 2016 - Oct. 28, 2016

Sub-Scale Launch Motor Selection - - Nov. 30, 2016

Sub-Scale Launch - - Dec. 11, 2016

Design Report - Dec. 2, 2016 Nov. 29, 2016

Design Presentation - Dec. 5, 2016 Dec. 2, 2016

Motor Mount Design/ FEA Jan. 13, 2017 - Nov. 30, 2016

All Structural elements FEA Jan. 13, 2017 - Nov. 30, 2016

CDR Report Jan. 13, 2017 - Dec. 9, 2016

CDR Flysheet Jan. 13, 2017 - Dec. 9, 2016

CDR Presentation Jan. 13, 2017 - Jan. 11, 2017

Full Scale Launch - - Feb. 12, 2017

FRR Report Mar. 6, 2017 - Mar. 1, 2017

FRR Flysheet Mar. 6, 2017 - Mar. 1, 2017

FRR Presentation Mar. 6, 2017 - Mar. 3, 2017

Competition Apr. 5, 2017 - Apr. 5, 2017

LRR Report Apr. 6, 2017 - Apr. 3, 2017

UE Final Report - Apr. 17, 2017 Apr. 12, 2017

UE Final Presentation - Apr. 20, 2017 Apr. 17, 2017

PLAR Report Apr. 24, 2017 - Apr. 21, 2017

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Project ACE Gantt Chart Period Highlight: 27

T/M RESPONSIBLEPERIODS

1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 33 34 35

Proposal David 1 4 1 4100%

Preliminary Design Report David 6 4 6 4100%

PDR Presentation David 8 2 8 2100%

Interim Design Report David 12 3 13 2100%

Critical Design Report David 11 5 10 8100%

CDR Presentation David 15 5 15 5100%

Flight Readiness Report David 23 4 24 3100%

FRR Presentation David 26 2 26 2100%

Project Final Report David 31 20%

Launch Readiness Review David 29 40%

Post Launch Assesment David 33 20%

Budget Creation David 1 1 1 5100%

Website Creation Bryan 1 3 2 2100%

Motor Type Selection Andrew G 1 3 1 3100%

Motor Mount Design Andrew G 1 5 1 5100%

Rocksim Model Andrew G 3 18 3 4100%

Body Component Selection Torsten 1 6 1 6100%

3D Rocket Model Torsten 4 11 4 6100%

CFD Model Torsten 15 6 13 2100%

Payload A Design Justin 1 9 1 9100%

Payload B Design Braden 1 11 1 11100%

Data Acquisition Design David 3 6 3 2100%

Data Transmission Design David 3 6 3 2100%

Design of Recovery System Andrew S 1 9 1 9100%

Design Tracking System Andrew S 9 4 9 4100%

Design Education Activity Bryan 1 8 1 12100%

Propulsion Construction Andrew G 10 6 12 4100%

Body Construction Torsten 11 11 10 13100%

Payload A Construction Justin 9 14 9 14100%

Payload B Construction Braden 12 11 12 12100%

Recovery System Construction Andrew S 9 12 9 13100%

Data Systems Construction David 8 13 9 13100%

Scale Model Construction Torsten 12 3 11 4100%

Scale Model Test Team 14 2 14 2100%

Bulkhead Testing Rakan 22 3 24 2100%

Payload Testing Braden 14 9 20 6100%

Parachute Testing Andrew S 23 6 25 1100%

Wind Tunnel Testing Feras 23 9 25 330%

Recovery Testing Andrew S 19 7 22 4100%

Educational Engagement Bryan Ongoing Ongoing100%

Preparation for Competition David 31 10%

Competition David 32 10%

Rep

ort

ing

Des

ign

Co

nst

ruct

ion

Tes

tin

g

% (Unplanned)

ACTIVITY PLAN STARTPLAN

DURATION

ACTUAL

START

ACTUAL

DURATIONPERCENT COMPLETE

Plan Duration Actual Start % (Planned) Actual (beyond plan)

(Week 1 ends September 4th, 2016)

Figure 52 - Gantt Chart

179 | P a g e

References

Autodesk. (2015, December 28). External Incompressible Flow. Retrieved from Autodesk

Knowledge Network: https://knowledge.autodesk.com/support/cfd/learn-

explore/caas/CloudHelp/cloudhelp/2014/ENU/SimCFD/files/GUID-4EED9E6E-A694-

4505-9502-8D9CC42A5EC2-htm.html

Center, G. C. (2016, 08 10). 2017 NASA's Student Launch. Retrieved 08 11, 2016, from NASA:

http://www.nasa.gov/sites/default/files/atoms/files/nsl_un_2017_web.pdf

Engineering Toolbox. (n.d.). U.S. Standard Atmosphere. Retrieved from Engineering Toolbox:

http://www.engineeringtoolbox.com/standard-atmosphere-d_604.html

G. Lengellé, J. D. (2004, January). Combustion of Solid Propellants. Research Scientists,

Energetics Department Office national détudes et de recherches aérospatiales (ONERA).

Lofton, J. (2016, November 29). Mechanical Engineering Professor. (T. Maier, Interviewer)

Michael J. Moran, H. N. (2014). Fundamentals of Engineering Thermodynamics. Hoboken: John

Wiley & Sons, Inc.

NASA. (n.d.). 2017 NASA Student Launch: Colleges, Universities, Non-Academic Handbook.

2017.

Niskanen, S. (2009). Development of an Open Source model rocket simulation software.

OpenRocket. Helsinki: HELSINKI UNIVERSITY OF TECHNOLOGY.

Ring, C. (2016, 9 27). Launch Crue. Retrieved from LaunchCrue.org:

https://www.launchcrue.org/

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Schmidt, D. P. (2016, October 15). Natural Frequency.

Weidong Cai, P. T. (2008). A MODEL OF AP/HTPB COMPOSITE PROPELLANT

COMBUSTION IN ROCKET-MOTOR ENVIRONMENTS. Taylor & Francis Group,

LLC.

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Appendix A – Machine Prints

Dimensioned Drawings

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Figure 53 – Aft Body Tube Drawing

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Figure 54 - Bow Body Tube Drawing

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Figure 55 - Fin Drawing

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Figure 56 - Motor Drawing

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Figure 57 - Nosecone Drawing

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Figure 58 - Launch Vehicle Drawing

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Figure 59 – Recovery bulkhead drawing

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Figure 60 - Payload Main bulkhead residing in Cylinder 2

Figure 61 - Payload assembly general dimensions

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Figure 62 - Recovery attachment bulkhead and hardware

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Figure 63 - Altimeter Mounting Plate Piece 1

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Figure 64 - Altimeter Mounting Plate Vertical 1

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Figure 65 - Metal O-Ring

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Figure 66 – Propulsion Section

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Figure 67 –Inner Tube

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Figure 68 - Centering Ring

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Figure 69 - Thrust Plate

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Figure 70 - Inner Cylinder

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Figure 71 - Payload Coupler

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Appendix B – OpenRocket Simulation

Sub-scale OpenRocket Inputs

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Appendix C – Best Fit Curve

OpenRocket Simulation Piecewise Regression

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Appendix D – OpenRocket Simulation

Inputs for OpenRocket Flight Simulation and Different Flight Configurations

Flight Configuration 1

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Flight 2 Configuration

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Flight 3 Configuration

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Appendix E – Payload Part Specification

Payload Part Specification Sheets

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Appendix F – Line Item Budget

Line Item Budget

Section Item Description Part Number Manufacturer Lead Time (days) Quantity Price (ea) Price (total) Date Ordered 20-02-290133 Amt. 20-01-240052 Amt. 20-01-240052 Amt. AFB

Nose Cone 5.5" FIBERGLASS 4:1 OGIVE NOSE CONE 20540 Apogee 1 84.95$ 84.95$ 17-Nov 84.95$

Body Tube 5.5" x 48" Carbon Fiber Airframe Wildman Rocketry 30 days 2 358.80$ 717.60$ 10-Nov 677.98$

Fins G10 FIBERGLASS SHEET 1/4" X 1 SQ FT 14154 Apogee 4 54.00$ 216.00$ 13-Jan - 65.35$

Nose Cone Threads Adhesive Mount Nut 98007A013 McMaster 10 $1.44 14.44$ 10-Nov 14.44$

Nose Cone Bolts Stainless Steel Button-Head Socket Cap Screws 98164A134 McMaster 50 0.13$ 6.54$ 10-Nov 6.54$

Rail Buttons LARGE AIRFOILED RAIL BUTTONS (FITS 1.5" RAIL - 1515) 13069 Apogee 2 10.00$ 20.00$ 17-Nov 20.00$

Apogee Shipping 11/17 Order 1 41.63$ 41.63$ 17-Nov 41.63$

Extended Allen Wrenches Amazon 52.09$ 52.09$ 30-Jan 52.09$

1,153.25$ 845.54$ 117.44$

Motor AeroTech L850W 7525S AeroTech 1 1,420.00$ 1,420.00$ 1-Dec 1,526.75$

Retaining System Aero Pack 75mm Retainer - P 24055 Apogee 1 47.08$ 47.08$ 28-Nov 103.90$

Epoxy G5000 Rocketpoxy 2-pint package 30511 Apogee 2 38.25$ 76.50$ 10/18/2016 38.25$ 50.51$

Motor Mount 75mm Blue Tube 48" 10504 Apogee 1 29.95$ 29.95$ 17-Nov 29.95$

Motor Reloads AeroTech L850W Refuels 12850P AeroTech 2 199.99$ 399.98$ 1-Dec 252.98$

Centering Rings and Bulkheads .250" Aluminum Plate 6061-T651 2x4 P314T6 Metals4uOnline 7 1 181.50$ 181.50$ 18-Oct 181.50$

Lubricant Synthetic Grease 3 Pack - Amazon 1 23.08$ 23.08$ 10-Jan 23.08$

Igniters Fat Boy Starters 89885 Apogee 2 14.11$ 28.22$ 22-Nov 28.22$

2,206.31$ 381.82$ 1,853.32$

U-Bolts w/mounting plates for use with aluminum bulkhead (pack of 5) 3043T68 McMaster 1 5.89$ 5.89$ 10-Nov 5.89$

Electronics bay coupler 5.5" OD, bulkheads, rails 10526 Apogee 1 56.95$ 56.95$ 17-Nov 56.95$

Igniter terminal block for easy igniter replacement 9191 Apogee 2 3.41$ 6.82$ 10/18/2016 6.82$

Crimp Connector pack of 5 - Lowe's 1 4.98$ 4.98$ 10-Nov 4.98$

Ejection well 2-pack PVC wells for black powder 3068 Apogee 2 3.15$ 6.30$ 10/18/2016 6.30$

Parachute Protector 18" Nomex flameproof cloth 29314 Apogee 2 10.49$ 20.98$ 10/18/2016 20.98$

Tubular Nylon Recovery Harness 30351 Onebadhawk 2 38.00$ 76.00$ 10-Nov 84.00$

Shock Cord Protector 30" flameproof sheath 29300 Apogee 2 12.95$ 25.90$ 10/18/2016 25.90$

Rotary Switch lockable switch 9128 Apogee 2 9.93$ 19.86$ 10/18/2016 19.86$ 24.74$

Shear Pins Nylon, threaded (10 pack) 29615 Apogee 10 3.10$ 31.00$ 10/18/2016 31.00$

Quick Links link eyebolts, chutes, and cord - Lowe's 6 2.48$ 14.88$ -

24" Drogue Chute 24" Classic Elliptical Chute 29163 Apogee 1 63.70$ 63.70$ 17-Nov 63.70$

96" Main Chute Torroidal, 2.2Cd, Ripstop Nylon 29185 Apogee 1 346.53$ 346.53$ 17-Nov 346.53$

Stratologger CF Main & Backup 9104 Apogee 2 58.80$ 117.60$ 10/18/2016 117.60$ 123.73$

Ejection Charge Starters QBECS QuickBurst 30 1.25$ 37.50$ 22-Nov 45.50$

Parachute Slider slows parachute deployment Giant Leap Rocketry 1 13.22$ 13.22$ 22-Nov 22.34$

Black Powder - Gun Store 1 20.00$ 20.00$ 6-Dec 111.38$

9 Volt Battery Pack of 4 - Lowe's 1 12.47$ 12.47$ 15-Nov 48.43$

Zip Ties Pack of 100 Lowe's 1 4.48$ 4.48$ 15-Nov -

USB Data Transfer Kit PerfectFlite 1 22.46$ 22.46$ 22-Nov 22.86$

16 Gauge Wire - Lowe's 1 5.41$ 5.41$ 15-Nov -

912.93$ 929.64$ 259.85$

Atlus Metrum TeleMega From csrocketry.com Atlus Metrum 21 1 406.10$ 406.10$ 13-Oct 411.10$

Starter Pack From csrocketry.com Atlus Metrum 0 1 100.00$ 100.00$ 13-Oct 100.00$

Arrow 440-3 Yagi Antenna get from link in start pack page Yagi 0 1 50.00$ 50.00$ 18-Oct 54.00$

SMA to BNC adapter From csrocketry.com Atlus Metrum 0 1 10.00$ 10.00$ 18-Oct 19.00$

Washers McMaster, For Spacing & Damping 90133A005 McMaster 3 1 6.81$ 6.81$ 10-Nov 6.81$

O-Ring Bolts 10-24, 9/16in 91864A091 McMaster 3 1 $10.69 10.69$ 10-Nov 10.69$

Altimiter Bolts 5-40, 5/8in 91251A130 McMaster 3 1 $8.98 8.98$ 10-Nov 8.98$

Studs for Ballast .25 x 40, 1 in long 98750A011 McMaster 3 4 $1.07 4.28$ 10-Nov 4.28$

596.86$ 614.86$ -$

1/12 McMaster Order -$ 23.39$

Wire Rope Isolators First & Second Order 173.40$ 173.40$ 2-Dec 187.16$ 173.40$

Blue Tube (Testing) 5.5" x 48" Carbon Fiber Airframe 10506 Apogee - 1 56.95$ 56.95$ 17-Nov 56.95$

Outer Cylinder (Coupler) 5.36" OD, 5.217" ID Blue Tube 13106 Apogee 1 18.95$ 18.95$ 17-Nov 18.95$

Fastening Nuts For 3/8" x 16 Bolt, 1/4" Height 91813A190 McMaster 1 11.08$ 11.08$ 10-Nov 11.08$

Fastening Bolts 3/8" x 16 x 1" 91251A621 McMaster 1 8.62$ 8.62$ 14-Oct 23.39$

Base Washer 0.5" ID 1.25" OD 98026A114 McMaster 3 7.47$ 22.41$ 10-Nov 22.41$

Studs 3/8" x 1" Length 95475A624 McMaster 1 9.41$ 9.41$ 10-Nov 9.41$

Recovery Bolts 3/8" x 1.25" Length 91251A626 McMaster 1 9.27$ 9.27$ 10-Nov 9.27$

Recovery Nuts 3/8" Flanged 96282A103 McMaster 1 6.98$ 6.98$ 10-Nov 6.98$

Spacing Pipe 5.25" OD and 4.75" OD 8486K954 McMaster 1 57.46$ 57.46$ 10-Nov 57.46$

Springs Part Number 866 866 Century Spring Corp. 5 12.60$ 63.00$ 3-Nov 62.97$

Payload 2 Materials Apogee, Mcmaster, Century Spring 1 122.57$ 122.57$ 3-Feb 122.57$

McMaster Shipping 11/10 Order - - 1 6.78$ 6.78$ 10-Nov 6.78$

566.88$ 472.81$ 319.36$

Binder Staples Order - 4 1" & 3 1.5" - Staples - 1 55.22$ 55.22$ 12-Jan 55.22$

Misc. 1-19 19.61$

55.22$ -$ 74.83$

RockSim Temporary, 1 Seat License 1123 Apogee 0 1 20.00$ 20.00$ 20.00$

Jan. Amazon Order 2x Tap & Car Charger Amazon 1 24.97$ 24.97$ 24.97$

Shirts Notable Sponsors 3 43.33$ 130.00$ 120.00$

Hotel (Group A) Apr. 5 - 9, 2/Room, Avg. $120/night 10 People - - 4 360.00$ 1,440.00$ 7-Dec 1,661.28$

Hotel (Group B) Two Nights, 2/Room, Avg $120/night 4 People - - 2 240.00$ 480.00$

Fuel Reiumbursement 540mi/15mpg*$2.50/ga 5 Vehicles - - 5 90.00$ 450.00$

Shirt Cost For non professors 15 10.00$ 150.00$ 420.00$ (120.00)$

Embroidery 18 Shirts 20 18 105.93$

Memphis Hotel 1 Night 242.21$ 242.21$ 12/10/2016 242.21$

Louisville Re-Load L850W Reload 1 210.99$ 210.99$ 210.99$

Memphis Fuel 91.03$

2,694.97$ 560.00$ 10.90$ 2,205.51$

1515 Rail 1515 Extruded Al., 145" 16U252 Grainger 2 1 140.71$ 140.71$ 17-Oct 140.71$

Rail Bracket 90 Degree 5 Hole Bracket 47065T271 McMaster 2 4 9.74$ 38.96$ 17-Oct 38.96$

Bolts M10 x 20 x 1.5 91290A516 McMaster 2 1 6.41$ 6.41$ 17-Oct 6.41$

Shipping (McMaster) 11.51$ 17-Oct 11.51$

197.59$ 197.59$ -$

Body Tube 3" CARBON FIBER TUBING 60 INCHES LONG CFT3.0-60 Wildman 30 days 1 218.50$ 218.50$ 10/19/2016 243.00$

Nose Cone 3" FIBERGLASS 4:1 OGIVE NOSE CONE 20520 Apogee 1 30.95$ 30.95$ 10/18/2016 30.95$

Fins G-10 Fiberglass Sheet 0.125" (1/8") 12" x 24" Giant Leap 1 95.00$ 95.00$ 10/18/2016 95.00$

Rail Buttons LARGE AIRFOILED RAIL BUTTONS (FITS 1.5" RAIL - 1515) 13069 Apogee 2 10.00$ 20.00$ 10/18/2016 20.00$

Motor I435T 3836SC AeroTech 1 149.99$ 149.99$ 21-Oct 168.20$

Motor Reload I435T Reloads zero94314 AeroTech 2 54.99$ 109.98$ 21-Oct 87.98$

InnerTube 38mm BlueTube 10501 Apogee 1 16.49$ 16.49$ 10/18/2016 16.49$

Centering Rings/ Bulkhead Same as full scale/ use same sheet P314T6 Metal Depot - - 18-Oct -$

75mm Electronics Bay 10524 Apogee 1 39.93$ 39.93$ 10/18/2016 39.93$

48" Main Chute 29167 Apogee 1 126.85$ 126.85$ 10/18/2016 126.85$

18" Drogue Chute 29162 Apogee 1 56.90$ 56.90$ 10/18/2016 56.90$

20' Tubular Nylon Recovery Harness $5 shipping OneBadHawk 2 18.00$ 36.00$ 10/31/2016 41.00$

Eyebolts Lowe's 2 1.96$ 3.92$ 11/1/2016 4.00$

Misc mounting hardware Lowe's 1 10.00$ 10.00$ 11/2/2016 10.00$

Subscale Shipping 38.95$ 38.95$

Retention System AERO PACK 38MM RETAINER - P 24063 Apogee 1 26.75$ 26.75$ 11/28/2016 - -

Igniters Slim Gem Starters 89884 Apogee 1 13.80$ 13.80$ 11/22/2016 13.80$

Total 993.05$ -$

Total 994.01$ 4,995.31$ 2,635.70$ 2,205.51$

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219 | P a g e

Appendix G – Task Breakdown

Task Breakdown

Task* Responsible Comments

Person

1 Project Management David - -

   1.1 Proposal (Report) / Research David Aug. 15, 2016 Aug. 15, 2016 Sept. 6, 2016 Aug. 20, 2017 Complete

1.1.1 Create Standards for Proposal David May. 25, 2016 May. 25, 2016 Jun. 1, 2016 May. 5, 2016 Complete

      1.1.2 Write Proposal David Sept. 1, 2016 Sept. 1, 2016 Sept. 27, 2016 Sept. 26, 2017 Complete

1.2 Preliminary Design Review (Report) David - - -

1.2.1 Create Standards for Preliminary Design Review David Oct. 1, 2016 Oct. 1, 2016 Oct. 5, 2016 Oct. 5, 2016 Complete

1.2.2 Write Preliminary Design Review David Oct. 5, 2016 Oct. 5, 2016 Oct. 26, 2016 Oct. 26, 2016 Complete

1.3 Critical Design Review (Report) David - - -

1.3.1 Create Standards for Critical Design Review David Oct. 28, 2016 Oct. 28, 2016 Nov. 2, 2016 Nov. 2, 2016 Complete

1.3.2 Write Critical Design Review David Nov. 2, 2016 Nov. 2, 2016 Dec. 9, 2016 Dec. 9, 2016 Complete

1.4 Flight Readiness Review (Report) David - - -

1.4.1 Create Standards for Flight Readiness Review David Jan. 1, 2017 Jan. 1, 2017 Jan. 18, 2017 Jan. 18, 2017 Complete

1.4.2 Compile Flight Readiness Review David Feb. 1, 2017 Mar. 1, 2017

   1.5 Launch Readiness Review David - - -

1.5.1 Create Standards for Launch Readiness Review David Feb. 28, 2017 Mar. 3, 2017

1.5.2 Compile Lanch Readiness Review David Mar. 15, 2017 Apr. 3, 2017

1.6 Post - Launch Assesment (Report) David - - -

       1.6.1 Create Standards for Post Launch Assesment David Apr. 10, 2017 Apr. 12, 2017

       1.6.2 Compile Post Launch Assesment David Apr. 14, 2017 Apr. 21, 2017

1.7 Preliminary Design Review (Presentation) David - - -

1.7.1 Create Preliminary Design Review Presentation David Oct. 20, 2016 Oct. 20, 2016 Oct. 28, 2016 Oct. 28, 2016 Complete

1.7.2 Preliminary Design Review Practice David Oct. 28, 2016 Oct. 28, 2016 Oct. 28, 2016 Oct. 28, 2016 Complete

   1.8 Critical Design Review (Presentation) David - - -

       1.8.1 Create Critical Design Review Presentation David Jan. 1, 2017 Jan. 1, 2017 Jan. 11, 2017 Jan. 11, 2017 Complete

       1.8.2 Critical Design Review Practice David Jan. 11, 2017 Jan. 11, 2017 Jan. 11, 2017 Jan. 11, 2017 Complete

   1.9 Flight Readiness Review (Presentation) David - - -

       1.9.1 Create Flight Readiness Review Presentation David Feb. 25, 2017 Mar. 3, 2017

       1.9.2 Flight Readiness Review Practice David Mar. 3, 2017 Mar. 3, 2017

   1.10 Orchestrate Meetings David - - -

   1.11 Create Budget David Sept. 1, 2016 Sept. 1, 2016 Sept. 27, 2016 Sept. 26, 2016 Complete

       1.11.1 Budget Monitoring David - - -

   1.12 Create Schedule David May. 25, 2016 May. 25, 2016 Jun. 1, 2016 Aug. 25, 2017 Complete

   1.13 Create Detailed Task Breakdown David May. 1, 2016 May. 1, 2016 Jun. 1, 2016 May. 1, 2016 Complete

   1.14 Integration of Subsections David - - -

   1.15 Create and Maintain Website Bryan Sept. 12, 2016 Sept. 12, 2016 28-Apr Sept. 16, 2016 Complete

   1.16 Travel for Testing & Competition David Feb. 1, 2017 Feb. 1, 2017 Mar. 1, 2017

       1.16.1 Local Rocket Meetings David - - -

   1.17 Meet Course Deliverables David - - -

   1.18 Purchasing David - - -

   1.19 Time Cards David - - -

        1.19.1 Time Card Format Creation David May. 1, 2016 May. 1, 2016 May. 16, 2016 May. 1, 2016 Complete

        1.19.2 Weekly Time Card Compiling - - -

   1.20 HAM Radio Liscence  Justin

   1.21 Meetings - -

        1.21.1 Meeting Planning David - -

   1.22 Recruiting David Aug. 25, 2016 Aug. 25, 2016 Sept. 9, 2016 Sept. 8, 2016 Complete

Start Date

Estimated ActualActualEstimated

End Date

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2 Propulsion Andrew

  2.1 Motor Type Selection (General, Proposal Level) Andrew

2.1.1 Motor Research Andrew 1-Jul 1-Jul Aug. 19, 2016 Aug. 19, 2016 Complete

       2.1.2 Motor Comparision Andrew 1-Jul 1-Jul Sept. 14, 2016 Sept. 13, 2016 Complete

2.1.3 Motor Elimination Andrew 1-Jul 1-Jul Sept. 14, 2016 Sept. 13, 2016 Complete

2.1.4 Caclulate projected Altitude Andrew - 1-Jul - - Complete

2.1.5 Select projected motor Andrew Sept. 14, 2016 Sept. 14, 2016 Sept. 15, 2016 Sept. 13, 2016 Complete

2.2 Mission Performance Predictions Andrew -

2.2.2 Simulated Thrust Curve Andrew Sept. 14, 2016 Sept. 14, 2016 Sept. 15, 2016 Sept. 13, 2016 Complete

2.3 Conceptual Model Creation Andrew -

2.3.1 Motor Mount Design Andrew Aug. 15, 2016 Aug. 15, 2016 Sep. 4, 2016 Sept. 19, 2016 Complete

2.3.1.1 Motor Fastening Design Andrew Aug. 15, 2016 Aug. 15, 2016 Aug. 28, 2016 Sept. 19, 2016 Complete

2.3.1.2 Motor Placement Andrew Aug. 15, 2016 Aug. 15, 2016 Sep. 4, 2016 Sept. 4, 2016 Complete

2.3.1.3 Redesign Andrew Sept. 5, 2016 Sept. 1, 2016 Nov. 30, 2016 Oct. 5, 2016 Complete

2.3.2 Rear Aerodynamics Design Andrew

           2.3.2.1 Collaboration with Aerodynamics Andrew Aug. 15, 2016 Aug. 15, 2016 Nov. 30, 2016 Nov. 30, 2016 Complete

2.3.4 Ignition Design Andrew

2.3.4.1 Ignition Research Andrew Sept. 15, 2016 Sept. 15, 2016 Sept. 21, 2016 Sept. 19, 2016 Complete

2.3.4.2 Ignition Placement Andrew Sept. 15, 2016 Sept. 15, 2016 Sept. 21, 2016 Sept. 19, 2016 Complete

2.3.4.3 Ignition Fastening Design Andrew Sept. 15, 2016 Sept. 15, 2016 Sept. 21, 2016 Sept. 19, 2016 Complete

2.3.4.4 Ignition Safety Interlock Design Andrew Sept. 15, 2016 Sept. 15, 2016 Sept. 21, 2016 Sept. 19, 2016 Complete

2.3.4.5 Igniter Installation Hatch Design Andrew Sept. 15, 2016 Sept. 15, 2016 Sept. 21, 2016 Sept. 19, 2016 Complete

2.3.4.6 Launch Switch w/ Returning to "off" Position Andrew Sept. 15, 2016 Sept. 15, 2016 Sept. 21, 2016 Sept. 19, 2016 Complete

2.3.4.4 Redesign Andrew Sept. 15, 2016 Sept. 15, 2016 Nov. 30, 2016 Sept. 19, 2016 Complete

2.4 Rocksim Modeling Andrew -

2.4.1 Model Rocket with Motor w/ Different Weights Andrew Aug. 15, 2016 Aug. 15, 2016 Jan. 15, 2017 Oct. 5, 2016 Complete

2.4.1.1 Simulation 1 Andrew Aug. 30, 2016 Aug. 15, 2016 Sept. 14, 2016 Sept. 4, 2016 Complete

2.4.1.2 Discussion with Other Sections Andrew 15-Sep Sept. 15, 2016 Sept. 21, 2016 Sept. 6, 2016 Complete

2.4.1.2 Resimulate Andrew Sept. 21, 2016 Sept. 15, 2016 Sept. 30, 2016 Sept. 19, 2016 Complete

2.4.2 Simulate Full Scale Model Andrew

2.4.2.1 Preliminary Motor Selection Simulation Andrew Aug. 15, 2016 Sept. 1, 2016 Sept. 14, 2016 Sept. 13, 2016 Complete

2.4.2.2 Preliminary Weighted Sections Simulation Andrew Aug. 15, 2016 Sept. 1, 2016 Sept. 14, 2016 Sept. 13, 2016 Complete

2.4.2.3 Redesign Andrew Sept. 14, 2016 Sept. 14, 2016 Sept. 21, 2016 Sept. 19, 2016 Complete

           2.4.2.4 Final Motor Selection Simulation Andrew Sept. 15, 2016 Sept. 14, 2016 Sept. 21, 2016 Sept. 13, 2016 Complete

2.4.2.5 Second Weighted Section Simulation Andrew Sept. 21, 2016 Sept. 15, 2016 Sept. 25, 2016 Sept. 19, 2016 Complete

2.4.2.6 Redesign 2 Andrew Sept. 25, 2016 Sept. 15, 2016 Sept. 29, 2016 Sept. 19, 2016 Complete

2.4.2.7 Final Rocket Simulation Andrew Sept. 29, 2016 Sept. 27, 2016 Jan. 15, 2017 Jan. 15, 2016 Complete

2.4.3 Simulate Half Scale Model Andrew

2.4.3.1 Physical Similitude Calculations Andrew Sept. 14, 2016 Sept. 27, 2016 Nov. 30, 2016 Oct. 5, 2016 Complete

2.5 Preliminary Design Review Andrew

2.5.1 Baseline Motor Selection Andrew Sept. 15, 2016 Sept. 10, 2016 Sept. 16, 2016 Sept. 13, 2016 Complete

2.5.2 Thrust-Weight Ratio Andrew Sept. 15, 2016 Sept. 10, 2016 Sept. 16, 2016 Sept. 13, 2016 Complete

2.5.3 Rail Exit Veloctiy Andrew Sept. 15, 2016 Sept. 10, 2016 Sept. 16, 2016 Sept. 13, 2016 Complete

2.6 Critical Design Review David

2.6.1 Specify Motor Andrew Sept. 21, 2016 Sept. 10, 2016 Oct. 7, 2016 Sept. 13, 2016 Complete

2.6.2 Final Drawings Andrew Sept. 21, 2016 Sept. 21, 2016 Oct. 7, 2016 Oct. 7, 2016 Complete

2.6.3 Final Analysis and Model Results Andrew Sept. 29, 2016 Sept. 29, 2016 Dec. 5, 2016 Dec. 15 2016 Complete

2.6.4 Motor Mounts Andrew Sept. 5, 2016 Sept. 5, 2016 Nov. 30, 2016 Oct. 5, 2016 Complete

2.6.5 Altitude Predictions with Final Design Andrew Sept. 29, 2016 Sept. 27, 2016 Dec. 5, 2016 Sept. 19, 2016 Complete

2.6.6 Actual Motor Thrust Curve Andrew Sept. 29, 2016 Sept. 27, 2016 Dec. 5, 2016 Sept. 19, 2016 Complete

2.6.7 Show Scale Model Results Andrew Sept. 29, 2016 Sept. 29, 2016 14-Dec Dec. 15 2016 Complete

2.7 Critical Design Review Presentation David

        2.7.1 Final Motor Choice Andrew Sept. 15, 2016 Sept. 10, 2016 Oct. 7, 2016 Sept. 13, 2016 Complete

2.7.2 Rocket Flight Stability in Static Diagram Andrew Sept. 15, 2016 Sept. 15, 2016 Oct. 7, 2016 Oct. 7, 2016 Complete

2.7.3 Thrust-to-Weight ratio Andrew Sept. 15, 2016 Sept. 10, 2016 Oct. 7, 2016 Sept. 19, 2016 Complete

2.7.4 Rail Exit Velocity Andrew Sept. 15, 2016 Sept. 10, 2016 Oct. 7, 2016 Sept. 19, 2016 Complete

2.8 Flight Readiness Review Presentation David

2.8.1 Final Motor Choice/ description Andrew Sept. 15, 2016 Sept. 15, 2016 Oct. 7, 2016 Oct. 7, 2016 Complete

2.8.2 Key Design Features Andrew Sept. 21, 2016 Sept. 15, 2016 Nov. 30, 2016 Oct. 7, 2016 Complete

2.8.3 Rocket Flight Stability Andrew Sept. 15, 2016 Sept. 15, 2016 Oct. 7, 2016 Oct. 7, 2016 Complete

2.8.4 Launch Thrust-Weight Ratio Andrew Sept. 15, 2016 Sept. 15, 2016 Oct. 7, 2016 Oct. 7, 2016 Complete

2.8.5 Rail Exit Velocity Andrew Sept. 15, 2016 Sept. 15, 2016 Oct. 7, 2016 Oct. 7, 2016 Complete

2.9 Testing Andrew

2.9.1 Ignition Testing Andrew Dec. 11 2016 Dec. 5, 2016 Feb. 12, 2017 Dec. 5 2016 Complete

2.9.1.1 Switch Testing Andrew Dec. 11 2016 Dec. 5, 2016 Feb. 12, 2017 Dec. 5 2016 Complete

2.9.1.2 Fuel Igition Testing Andrew Dec. 11 2016 Dec. 5, 2016 Feb. 12, 2017 Dec. 5 2016 Complete

2.9.1.3 Ignition Mount Testing Andrew Dec. 11 2016 Dec. 5, 2016 Feb. 12, 2017 Dec. 5 2016 Complete

2.9.1.4 Ignition Safety Interlock Testing Andrew Dec. 11 2016 Dec. 5, 2016 Feb. 12, 2017 Dec. 5 2016 Complete

2.9.1.5 Misfire Testing Andrew Dec. 11 2016 Dec. 5, 2016 Feb. 12, 2017 Dec. 5 2016 Complete

2.9.2 Motor Testing Junior Dec. 11 2016 Not Applicable Feb. 12, 2017 Not Applicable

2.9.2.1 Impulse Testing Junior Dec. 11 2016 Not Applicable Feb. 12, 2017 Not Applicable

              2.9.2.1.1 Testing  Junior Dec. 11 2016 Not Applicable Feb. 12, 2017 Not Applicable

             2.9.2.1.2  Data Analysis Junior Dec. 11 2016 Not Applicable Feb. 12, 2017 Not Applicable

2.9.2.2 Thrust Testing Junior Dec. 11 2016 Not Applicable Feb. 12, 2017 Not Applicable

              2.9.2.2.1 Testing  Junior Dec. 11 2016 Not Applicable Feb. 12, 2017 Not Applicable

              2.9.2.2.2  Data Analysis Junior Dec. 11 2016 Not Applicable Feb. 12, 2017 Not Applicable

2.9.2.4 Motor Mount Testing Andrew Dec. 11 2016 Not Applicable Feb. 12, 2017 Not Applicable

2.9.3 FEA on Motor Mount Andrew

2.9.3.1 Vibration Analysis Andrew Sept. 19, 2016 Sept. 19, 2016 Dec. 1, 2016 Oct. 11, 2016 Complete

2.9.3.2 Combustion Analysis Andrew Sept. 19, 2016 Sept. 19, 2016 Dec. 1, 2016 Oct. 27, 2016 Complete

2.9.3.3 Modal Analysis Andrew Nov. 7, 2016 Oct. 24, 2016 Dec. 1, 2016 Oct. 24, 2016 Complete

2.9.3.4 Stiffness Analysis Andrew Sept. 19, 2016 Oct. 17, 2016 Dec. 1, 2016 Oct. 19, 2016 Complete

2.9.3.5 Impulse Analysis Andrew Sept. 19, 2016 Oct. 17, 2016 Dec. 1, 2016 Oct. 19, 2016 Complete

2.9.3.6 Shear Stress Calculations Andrew Sept. 29, 2016 Oct. 24, 2016 Dec. 1, 2016 Oct. 24, 2016 Complete

2.9.3.7 Shear Stress Analysis with FEA Andrew Nov. 7, 2016 Nov. 14, 2016 Dec. 1, 2016 Nov. 15, 2016 Complete

2.10 Construction

2.10.1 Centering Ring CNC Andrew Nov. 7, 2016 Nov. 07, 2016 Dec. 1, 2016 Nov. 21, 2016 Complete

2.10.2 Bulkhead CNC Andrew Nov. 7, 2016 Nov. 07, 2016 Dec. 1, 2016 Nov. 21, 2016 Complete

       2.10.3 Motor Mount Andrew Nov. 7, 2016 Nov. 07, 2016 Dec. 1, 2016 Nov. 28, 2016 Complete

221 | P a g e

3 Aerodynamics Torsten

3.1 3D Modeling - Entire Rocket Torsten 1-May 1-May Oct. 26, 2016 Oct. 26, 2016 Complete

       3.1.1 General, Proposal-Level Rocket Model & Component Selection Torsten 1-May 1-May Sep. 30, 2016 Sep. 30, 2016 Complete

       3.1.2 Integration of Subcomponent Models into 3D Model Torsten 1-Aug 1-Aug Oct. 26, 2016 Oct. 26, 2016 Complete

       3.1.3 1/2 Scale 3D Model Torsten 1-Nov 1-Nov Nov. 20, 2016 Nov. 20, 2016 Complete

       3.1.4.Wind Tunnel Scale 3D Model Torsten 1-Feb Jan. 25, 2017 Mar. 5, 2017

3.2 Fins, Body, Nose Cone Selection Torsten 1-Aug 1-Aug Oct. 9, 2016 Oct. 9, 2016 Complete

       3.2.1 Full Scale Selection Torsten 1-May 1-May Sep. 30, 2016 Sep. 30, 2016 Complete

3.2.2 1/2 Scale Selection Torsten 1-Nov 18-Oct Nov. 20, 2016 Sep. 30, 2016 Complete

3.2.3 Wind Tunnel Scale Selection Torsten 30-Mar Mar. 5, 2017

    3.3 Fins, Body, Nose Cone Construction Torsten

       3.3.1 Full Scale Construction Torsten Jan. 22, 2017

       3.3.2 1/2 Scale Construction Torsten 30-Nov 11-Nov Dec. 4, 2016 Dec. 4, 2016 Complete

       3.3.3 Wind Tunnel Scale Construction Torsten 12-Jan Apr. 2, 2017 Behind Schedule

3.4 Paint Torsten

3.4.1 Paint Effect on Drag Torsten 1-Aug 1-Aug Oct. 26, 2016 Oct. 26, 2016 Complete

3.4.2 Painting Torsten Not happening Not happening Not happening Not happening Complete

3.5 Determination of Center of Mass Torsten 1-Aug 1-Aug Jan. 22, 2017 Jan. 12 Complete

3.6 Determination of Center of Pressure Torsten 1-Aug 1-Aug Jan. 22, 2017 Jan. 12 Complete

3.7 Optimization of Center of Mass vs Center of Pressure Torsten 1-Aug 1-Aug Jan. 22, 2017 Jan. 12 Complete

3.8 CFX Modeling Torsten Jan. 15, 2016 30-Dec Complete

3.8.1 Full Scale Rocket Performance Torsten 1-May 1-May Dec.12, 2016 30-Dec Complete

3.8.2 1/2 Scale Rocket Performance Torsten 1-Nov N/A Nov. 20, 2016 N/A

3.8.3 Wind Tunnel Scale Performance Torsten 12-Jan Behind Schedule Mar. 5, 2017

3.9 Collaboration with Launch Pad for Guides Torsten 1-Nov 18-Oct Jan. 22, 2017 Jan. 12 Complete

3.10 Study Feasability of Real-Time Drag Changing Torsten 1-Aug 1-Aug Sep. 30, 2016 Sep. 30, 2016 Complete

3.11 Redesign of Rocket Body, Nosecone, Fins Torsten 1-Nov 18-Oct Jan. 22, 2017 Jan. 12 Complete

4 Payload A

4.1 Payload A Design Justin - -

4.1.1 Official Altimeter Justin Aug. 20, 2016 Aug. 20, 2016 Sept. 20, 2016 Sept. 1, 2016 Complete

4.1.2 Radio Frequency and GPS Tracking Justin Aug. 20, 2016 Aug. 20, 2016 Sept. 20, 2016 Sept. 1, 2016 Complete

4.1.3 Arming and Disarming Electronics Justin Aug. 20, 2016 Aug. 20, 2016 Sept. 20, 2016 Sept. 20, 2016 Complete

4.2 Payload A Construction Justin Nov. 1, 2016 Nov. 20, 2016

4.2.1 Official Altimeter Justin Nov. 1, 2016 Nov. 1, 2016 Nov. 20, 2016 Nov. 21, 2016 Complete

4.2.2 Radio Frequency and GPS Tracking Justin Nov. 1, 2016 Nov. 1, 2016 Nov. 20, 2016 Nov. 21, 2016 Complete

4.2.3 Arming and Disarming Electronics Justin Nov. 1, 2016 Nov. 1, 2016 Nov. 20, 2016 Nov. 21, 2016 Complete

4.3 Payload A Redesign Justin Nov. 10, 2016 Nov. 10, 2016 Nov. 20, 2016 Nov. 21, 2016 Complete

4.4 Integration with Data Collection System Justin Aug. 20, 2016 Aug. 20, 2016 Nov. 28, 2016 Nov. 28, 2016 Complete

4.5 Data Transmission Justin - -

4.5.1 Wireless Receiver Justin Aug. 20, 2016 Aug. 20, 2016 Nov. 1, 2016 Nov. 1, 2016 Complete

4.5.1.1 Design Ground Station Wireless Receiver Justin Aug. 20, 2016 Aug. 20, 2016 Nov. 1, 2016 Nov. 1, 2016 Complete

4.5.1.2 Construct Ground Station Wireless Reciever Justin Nov. 1, 2016 Nov. 1, 2016 Nov. 20, 2016 Nov. 21, 2016 Complete

4.5.2 Wireless Transmission Justin - - - -

4.5.1.1 Design Wireless Transmitter Justin Aug. 20, 2016 Aug. 20, 2016 Nov. 20, 2016 Nov. 21, 2016 Complete

4.5.1.2 Construct Wireless Transmitter Justin Nov. 1, 2016 Nov. 1, 2016 Nov. 20, 2016 Nov. 21, 2016 Complete

4.6 Create Test Plan & Test to Ensure Components in working order Justin Nov. 1, 2016 Nov. 1, 2016 Dec.12, 2016 Behind Schedule

4.7 Collaboration with Payload B over Motherboard Justin - - - -

4.8 Determine if Separation is Necessary Justin Aug. 20, 2016 Aug. 20, 2016 Sept. 20, 2016 Sept. 20, 2016 Complete

4.9 Ensure that all components can be subjected to rocket stresses Justin Nov. 1, 2016 Nov. 1, 2016 20-Jan 14-Nov Complete

4.10 Meetings/Reports Justin - -

5 Payload B Braden

5.1 Payload B Design (Fragile Material Housing) Braden

5.1.1 Design of Experiment Braden Aug. 1, 2016 Aug. 1, 2016 Sep. 5, 2016 Not completed Complete

5.1.2 Design of Experimental Apparatus Braden Sep. 5, 2016 Sep. 5, 2016 Ongoing and changing Sept. 19, 2016 Complete

5.1.3 Design of Mounting Braden Sep. 5, 2016 Sept. 15, 2016 Sep. 20, 2016 Sept. 20, 2016 Complete

5.2 Payload B Construction Braden

5.2.1 Construction of Experiment and housing Braden Oct. 1 1-Nov Ongoing and changing 1/17/2017 Complete

5.2.2 Construction of Mounting Braden Oct. 1 1-Nov Nov. 20, 2016 1/17/2017 Complete

5.3 Payload Testing and Experimentation Braden

5.3.1 Design Testing Plan Braden Sept. 10, 2016 Nov 1. 2016 Sep. 30, 2016 Nov. 28, 2016 Complete

5.3.2 Carry Out Testing Braden Oct. 10, 2016 30-Jan Dec. 4, 2016 15-Feb Complete

5.3.3 Data Analysis Braden Dec. 4, 2016 30-Jan Jan. 22, 2017 15-Feb Complete

5.3 Payload B Redesign Braden Jan. 22, 2017 Jan. 22, 2017 feb. 1, 2017 feb. 1, 2017 Complete

     5.7 Ensure that all components can be subjected to rocket stresses Braden Jan. 22, 2017 Jan. 22, 2017 feb. 1, 2017

5.7 Reports Braden

5.7.1 PDR Braden Sept. 15, 2016 Sept. 15, 2016 Sept. 19, 2016 Sept. 19, 2016 Complete

5.8 Meetings/Group Work Braden

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6 Recovery Stewart 9-Jan 9-Jan 3-Feb 25-Jan COMPLETE

6.1 Recovery System Design Stewart 15-Aug 15-Aug 30-Sep 30-Sep COMPLETE

          6.1.1 Recovery System Research Stewart 15-Aug 15-Aug 9-Sep 14-Oct COMPLETE

6.1.2 Recovery System Component Selection Stewart - - - -

6.1.2.1 Parachutes (Drogue & Main) Stewart 29-Aug 12-Sep 30-Sep 16-Oct COMPLETE

6.1.2.2 Altimeters Stewart 29-Aug 12-Sep 9-Sep 16-Oct COMPLETE

6.1.2.3 Shock cord and hardware Stewart 29-Aug 12-Sep 9-Sep 16-Oct COMPLETE

6.1.2.4 Ejection system Stewart 29-Aug 12-Sep 9-Sep 16-Oct COMPLETE

6.1.2.5 Bulkhead components Stewart 29-Aug 12-Sep 9-Sep 16-Oct COMPLETE

               6.1.2.6 Electronics Stewart 29-Aug 12-Sep 9-Sep 16-Oct COMPLETE

6.1.3 Bulkhead design Stewart 29-Aug 12-Sep 30-Sep 23-Oct COMPLETE

6.1.4 Circuit design & programming Stewart 29-Aug 19-Sep 30-Sep 16-Oct COMPLETE

6.1.5 Computer Modeling - - - -

           6.1.5.1 Parachute modeling Stewart 29-Aug 19-Sep 30-Sep 16-Oct COMPLETE

6.1.5.2 3D Assembly 29-Aug 12-Sep 30-Sep 23-Oct COMPLETE

6.1.6 Scaled model design Stewart - - - -

               6.1.6.1 Parachutes (Drogue & Main) Stewart 29-Aug 10-Oct 30-Sep 16-Oct COMPLETE

               6.1.6.2 Shock cord and hardware Stewart 29-Aug 10-Oct 9-Sep 16-Oct COMPLETE

               6.1.6.3 Bulkhead components Stewart 29-Aug 10-Oct 9-Sep 7-Nov COMPLETE

               6.1.6.4 Ejection system Stewart 29-Aug 10-Oct 9-Sep 16-Oct COMPLETE

6.2 Recovery System Construction Stewart 2-Dec

6.2.1 Bulkhead assembly Stewart 9-Jan 9-Jan 23-Jan 23-Jan COMPLETE

          6.2.2 Circuit assembly Stewart 7-Nov 14-Nov 23-Jan 23-Jan COMPLETE

6.2.3 Ejection system assembly Stewart 9-Jan 14-Nov 23-Jan 23-Jan COMPLETE

6.2.4 Full-system integration Stewart 9-Jan 14-Nov 2-Dec 2-Dec COMPLETE

6.2.5 Scaled model construction Stewart 31-Oct 14-Nov 2-Dec 2-Dec COMPLETE

6.3 Recovery System Testing Stewart

6.3.1 Parachute testing (multiple wind speeds) Stewart 5-Dec 22-Jan 3-Feb 3-Feb Completed

6.3.2 Ejection system testing Stewart 9-Jan 22-Jan 20-Jan 20-Jan Completed

6.3.3 Circuit and transmitter testing Stewart 9-Jan 30-Nov 20-Jan 20-Jan Completed

6.3.4 Full-system testing Stewart 23-Jan 20-Jan 3-Feb 3-Feb Completed

    6.4 Launch Pad David

6.4.1 Launch Pad Design David Aug. 29, 2016 Aug. 29, 2016 Sept. 30, 2016 Sept. 30, 2016 Completed

6.4.2 Launch Pad Material Aquisition David Sept. 30, 2016 Sept. 30, 2016 Oct. 10, 2016 Oct. 10, 2016 Completed

6.4.3 Launch Pad Fabrication David Oct. 20, 2016 Oct. 20, 2016 Oct. 25, 2016 Oct. 25, 2016 Completed

6.5 Obtain Launch License Stewart 4-Nov 4-Nov 4-Dec 14-Nov Completed

7 Testing Bryan

7.1 Oversee all Subsection Testing Bryan Dec. 12, 2016 Nov. 11, 2016 5-Apr

7.2 Manage Junior Level Testing Bryan Dec. 12, 2016 12-Dec 17-Mar Ongoing

7.3 1/2 Scale Testing Bryan - -

7.3.1 Design of 1/2 Scale Testing Experiments Bryan Sept. 30, 2016 Nov. 16, 2016 Dec. 2, 2016 Dec. 2, 2016 Completed

7.3.2 Construction and Conduction of 1/2 Scale Testing Experiments Bryan Dec. 2, 2016 Dec. 5, 2016 Dec. 7, 2016 Dec. 9, 2016 Completed

7.3.3 Assess CFX with Results Bryan Jan. 9, 2017 Jan. 14, 2017 Completed

7.4 Wind Tunnel Testing Bryan Feb. 5, 2017 Feb. 26, 2017 Behind Schedule

7.4.1 Assess CFX with Results Bryan 20-Mar 25-Mar Behind Schedule

7.5 Work with Subsections to Optomize Sections based on Testing Bryan Dec. 12, 2016 Dec. 7, 2016 25-Mar Completed

7.6 Modify Wind Tunnel for Scale Testing Bryan Feb. 26, 2017 17-Mar

7.7 Create Stand for Wind Tunnel Testing Bryan Jan. 31, 2017 Feb. 5, 2017 Behind Schedule

7.8 Assess Rocksim with Fullscale Data Bryan 17-Mar 25-Mar

7.9 Assess Rocksim with 1/2 Scale Test Bryan Dec. 2, 2016 Dec. 7, 2016 Dec. 9, 2016 Dec. 9, 2016 Completed

8 Safety Bryan

8.1 Create a Detailed Step-by-Step Launch Procedure Bryan Nov. 7, 2016 Oct. 3, 2016 Dec. 8, 2016 Oct. 21, 2016 Completed

8.1.1 Monitor Team Activities per NASA Handbook sec. 4.3 Bryan - -

8.1.2 Maintain all Safety Activities per NASA Bryan Aug. 29, 2016 Sept. 5, 2016 Dec. 2, 2016 Dec. 2, 2016 Completed

8.2 Designated Head of Safety Bryan - -

8.3 Creation of Safety Checklist Bryan Aug. 29, 2016 Sept. 5, 2016 Sept. 30, 2016 9/25/2016 Completed

8.4 Manage and Maintain MSDS Sheets Bryan - -

8.5 Manage and Maintain Hazard Analysis Documents Bryan - -

8.6 Manage and Maintain Failure Mode Analyses Bryan - -

9 Educational Engagement Bryan

9.1 Create and Orchestrate Educational Engagement Activity Bryan Sept. 1, 2016 Sept. 5, 2016 Feb. 15, 2017 Completed

9.2 Create Report for Educational Engagement Activity Bryan Nov. 7, 2016 Nov. 9, 2016 Feb. 15, 2017 Completed

9.3 Create Presentation for Educational Engagement Activity Bryan Nov. 14, 2016 Oct. 25, 2016 Feb. 15, 2017 Behind Schedule

9.4 Create Display for Educational Engagement Activity Bryan Nov. 14, 2016 Oct. 25, 2016 Feb. 15, 2017 Behind Schedule

223 | P a g e

Appendix H – Electrical Diagrams

Electrical Diagrams

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Appendix I – Payload Accelerometer Graphs

Payload Accelerometer Graphs

Figure 72 – 90 Degree Cotton Fill Large Bulb

-40000

-30000

-20000

-10000

0

10000

20000

30000

0 20 40 60 80 100 120

Acc

ele

rati

on

Time Step (s)

90 Degree Cotton Fill Large Bulb

Ax

Ay

Az

-

31727

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Figure 73 – 90 Degree Paper Fill Large Bulb

Figure 74 – 90 Degrees Packing Peanuts Large Bulb

-4000

-3000

-2000

-1000

0

1000

2000

3000

0 10 20 30 40 50 60

90 Degree Paper Fill Large Bulb

Ax

Ay

Az

-3434

-40000

-30000

-20000

-10000

0

10000

20000

30000

40000

0 20 40 60 80 100 120 140

90 Degrees Packing Peanuts Large Bulb

Ax

Ay

Az

32185

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Figure 75 – 90 Degrees Large Bulb Only

-3000

-2000

-1000

0

1000

2000

3000

0 1 2 3 4 5 6

90 Degrees Large Bulb Only (no fill)

Ax

Ay

Az

2526

-4000

-3000

-2000

-1000

0

1000

2000

3000

4000

0 5 10 15 20 25 30 35

90 Degrees DogBrag Fill Large Bulb

Ax

Ay

Az

3249

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Figure 76 – 90 Degrees DogBrag Fill Large Bulb

Figure 77 – 90 Degrees Base Value

-25000

-20000

-15000

-10000

-5000

0

5000

10000

15000

0 1 2 3 4 5 6 7 8 9

90 Degrees Base Value

Ax

Ay

Az

-18895

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Appendix J – Wind Tunnel Uncertainty

Wind Tunnel Uncertainty

Sample calculation of drag coefficient’s uncertainty:

Assuming density (𝜌) is constant.

𝐶𝐷 =𝜖∗𝑤∗𝑡2

6∗𝐿∗𝜌∗ 𝑈2

2∗𝐴𝑐

𝜕𝐶𝐷

𝜕𝜖=

𝐸∗𝑤∗𝑡2

6∗𝐿∗𝜌∗ 𝑈2

2∗𝐴𝑐

𝜕𝐶𝐷

𝜕𝐸=

𝜖∗𝑤∗𝑡2

6∗𝐿∗𝜌∗ 𝑈2

2∗𝐴𝑐

𝜕𝐶𝐷

𝜕𝑡=

2∗𝐸∗𝑤∗𝑡

6∗𝐿∗𝜌∗ 𝑈2

2∗𝐴𝑐

𝜕𝐶𝐷

𝜕𝐿= −(

𝜖∗𝐸∗𝑤∗𝑡2

6∗𝐿2∗𝜌∗ 𝑈2

2∗𝐴𝑐

)

𝜕𝐶𝐷

𝜕𝑈= −(

2 ∗ 𝜖 ∗ 𝐸 ∗ 𝑤 ∗ 𝑡2

6 ∗ 𝐿 ∗ 𝜌 ∗ 𝑈2

3

∗ 𝐴𝑐

)

𝜕𝐶𝐷

𝜕𝑈= −(

𝜖 ∗ 𝐸 ∗ 𝑤 ∗ 𝑡2

6 ∗ 𝐿 ∗ 𝜌 ∗ 𝑈2

2 ∗ 𝐴𝑐2

)

For velocity:

229 | P a g e

𝜕𝑈 𝜕∆𝑃

(𝑙𝑏𝑓 𝑓𝑡2

)⁄

= √2

0.00226 ∆𝑃(𝑙𝑏𝑓 𝑓𝑡

2)⁄

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Appendix K – MTS Tensile Test Procedure

1. Open the water inlet valve on the west wall of the lab.

2. Turn the large knob on HPU to the ON position.

3. Hit the reset button on the HPU.

4. Turn on the MTS control box, using the white power switch.

5. Turn on the computer.

6. When prompted to “Log On” the password is “admin” and then hit the OK button.

7. Double click the MTS station manager shortcut icon on the computer desktop.

8. From the Open Station Dialog Box, choose basic configuration.cfg as the configuration

file.

9. Reset the software interlock (red colored flag) from the Station Manager.

10. Check the exclusive control check box

11. From the Station Manager step up the power on the HPS. Start with low power and then

go to high power.

12. From the Station Manager step up the power on the HSM. Start with low power and then

go to high power.

13. Place the assembly in the grips and use the manual control knob to raise the bottom

platform.

14. Double click multipurpose elite on desktop.

15. Go to custom templates and double click it.

16. Double click NASA team test.

17. Click new test run.

18. Type the specimen name: we used rocket bulkhead.

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19. Then hit ok.

20. Close quote-custom window when it pops up.

21. Hit the green play button to start the test

22. Then the test is going to run. As the test runs, you should get a screen showing the load

and displacement curve. The test will stop automatically when the force level drops by

about 25% of the maximum.

23. Print the force versus displacement graph.

24. To find the data that was recorded, go to ME 330 lab data files. Under the specimens’

folder will be a folder with your specimen name. Use the excel file to obtain the data of

your specimen.

To Shut down the MTS machine:

25. From the station manager or the RSC, step the HSM down starting at the low position

and then to off.

26. From the station manager or the RSC, step the HPS down.

27. Close the station manager.

28. Turn off the controller by turning the switch in the back to off.

29. Shutdown the computer by using the shutdown option from the start button.

30. When the computer system is ready to be turned off hit the power button on the front

panel of the CPU.

31. Turn the large red knob on the HPU to the off position. Shut off the water to the HPU by

turning the yellow bale handle to the horizontal position.