University of Evansville
Student Launch
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
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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
퐴푃 Ammonium Perchlorate Composite
퐵퐻 Bulkhead
퐶푝 Specific Heat with constant pressure [kJ/kmol-K]
퐶푅 Centering Rings
퐸 modulus of elasticity [psi]
퐹 Force [lbf]
푓 operational frequency [Hz]
푓푛 natural frequency [Hz]
퐹. 푂. 푆 Factor of Safety
ℎ enthalpy [kJ/kmol] Combustion Analysis Section
ℎ thickness [in]
표 ℎ푓 Enthalpy of Formation [kJ/kmol]
퐼 second moment of inertia [𝑖푛4]
푘′ stiffness [psi]
KE Kinetic Energy
퐿 length [in]
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푀 moment [lbf-in]
푚 mass [lbm, kg] kg will be specified in the equation otherwise it is lbm
푛 moles [kmol]
푃 pressure [kPa]
푄 heat [kJ]
푅 radius [in]
푅 Gas Constant [=8.314 kJ/kmol-K] Combustion Analysis Section
푟 frequency ratio
푆퐴 surface area [𝑖푛2]
푇 temperature [K]
푡 time [s]
푉 volume [𝑖푛3]
푣 velocity of the rocket at burnout [m/s]
푣푓 ground impact velocity [ft/s]
푣푑 descent rate [ft/s]
휔 circular natural frequency [rad/sec]
푤 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.
3 A inch hole was drilled in the furthest aft centering ring. The hole allows the furthest aft rail 8 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 2 | P a g e 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
4 | P a g e 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. 5 | P a g e
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
6 | P a g e 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 Severity Likelihood Section Success Criteria Explanation of Failure of Failure A failure of the airframe during The airframe, nose cone, flight could cause a complete and fins should remain failure in the launch vehicle’s intact for the duration of Significant Low flight ability. However, the use the flight. of carbon fiber mitigates this risk
to a very low likelihood.
Reusability of parts is not The airframe, nose cone, detrimental to the project; new and fins should be parts can be purchased. The most reusable for any Minor Medium likely part to fail is a fin upon following flights. recovery, thus warranting a Aerodynamics medium likelihood of failure. The airframe and nose Damage to internal components cone should protect all can be detrimental to the launch internal components vehicle’s ability to deploy the Major Low from damage from recovery system. However, a external sources. carbon fiber airframe mitigates this risk to a very low likelihood.
Motor variations and launch day The vehicle should conditions both contribute to attain an apogee apogee variations from the full- between 5,125 feet and Medium Low scale test. However, the team’s
Propulsion 5,375 feet. allocated window should
account for these. 9 | P a g e
Severity Likelihood Section Success Criteria Explanation of Failure of Failure This is a requirement from The vehicle should NASA. The vehicle has not been remain below Mach 1. designed to withstand transonic Significant Low forces. The anticipated Mach number is 0.56. The motor mount should Motor mount/retention failure withstand propulsion could cause a poor flight, no forces and remain flight, or safety hazard. This is Major Low reusable for any mitigated by using aluminum following flights. and high strength epoxy.
If the drogue parachute does not deploy at apogee, the main The drogue parachute is parachute will deploy at high successfully deployed at Major Low velocity. This could result in apogee. damage to the parachute or airframe. If the main parachute does not The main parachute is deploy, the launch vehicle will successfully deployed at
descend under only the drogue Major Low 750 feet. parachute. This would result in
excessive ground impact speed. The drogue parachute A partially-inflated parachute is Recovery and main parachute much less effective at slowing inflate successfully the launch vehicle during its Major Low following ejection. descent. This could result in excessive ground impact speed. The maximum kinetic energy of any Excessive kinetic energy on independent section of landing could result in damage Major Low the rocket is less than 75 to the fragile payload or ft-lbf at landing. airframe.
The data sent from the electronic payload should be able to be If the data is not received received remotely remotely after the flight, the Low Low
Electronic Electronic during and after the team will not be scored. Payload vehicle’s flight.
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Severity Likelihood Section Success Criteria Explanation of Failure of Failure The electronic payload should withstand flight The altimeter notwithstanding forces and remain the forces will prevent the Low Low reusable for any altimeter from being reused, and following flights. the team cannot be scored.
The electronic payload Low should accurately If apogee is not detected
determine the apogee of accurately, it will affect the score Low the rocket. of the team.
The fragile object(s) should remain undamaged. To properly reduce the risk of The force felt by the damage to any and all unknown payload should be fragile material, the desired reduced by 50% for reduction of force felt by the Major Medium each of the areas of payload should be reduced by a interest: takeoff (thrust minimum of 50 percent for the curve, parachute most extreme forces exerted deployment, and throughout flight. landing.)
Main Payload Main The Acceleration felt by the payload should be To reduce the force at the reduced by 35% for maximum and minimum points each of the areas of of spring displacement, total Significant Low interest: (thrust curve, acceleration of the payload parachute deployment, should be reduced by a and landing.) minimum of 35 percent.
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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1 Figure 3. The body tube was then filed smooth using a file. The next step was drilling three 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
3 inches long, the fin tab has a height of 1.2 inches tall, and the fin tab length is five inches. inch 8 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 14 | P a g e
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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Outer Area Epoxy Location
Inner Area Epoxy Location
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
16 | P a g e 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
17 | P a g e 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.
EJECTION WELL U-BOLT
MOLEX TERMINAL CONNECTOR BLOCK
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 18 | P a g e 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.
COUPLING TUBE
ALL THREAD MOLEX ROD CONNECTOR
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.
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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 20 | P a g e
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 24 | P a g e
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
30 | P a g e 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 31 | P a g e 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
Launch Day Simulated Simulation Quick Links Ballast (lb) Conditions 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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6000
5000
4000
3000 Altitude(ft) 2000
1000
0 0 20 40 60 80 100 120 140 Time (s)
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
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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.
6000
5000
4000
3000 Flight 1
Flight 2 Altitude(ft) Flight 3 2000
1000
0 0 20 40 60 80 100 120 140 Time (s)
Figure 26 - Simulated Flight Configurations
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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 Velocity at Rail Exit
(s) (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 Predicted Altitude
(mph) (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)
41 | P a g e 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.
6000
5000 Actual Altitude (ft) 4000
3000 OpenRocket Altitude (ft) Altitude(ft) 2000
1000
0 0 20 40 60 80 100 120 140 160 Time (s)
Figure 28 - Flight 1 Actual vs OpenRocket Data
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6000
5000
4000 Actual Altitude (ft)
3000
OpenRocket Altitude (ft) Altitude(ft) 2000
1000
0 0 20 40 60 80 100 120 140 Time (s)
Figure 29 - Flight 2 Actual vs OpenRocket Data
6000
5000 Actual Altitude (ft) OpenRocket Altitude (ft)
4000
3000 Altitude(ft) 2000
1000
0 0 50 100 150 200 250 300 350 400 Time (s)
Figure 30 - Flight 3 Actual vs OpenRocket Data
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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:
(퐶 −퐶 ) 푆푡푎푏𝑖푙𝑖푡푦 = 푝 푔 (1) 퐷
In this equation, 퐶푝 is the Center of Pressure, 퐶푔 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
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Center of Pressure is a value determined by simulation from Open Rocket. 퐶푔 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 퐶푔 for each flight configuration was determined by hanging the rocket by a rope and balancing. 퐶푔 for each flight configuration is located at the balance point. This data can be found in Table 7 (The 퐶푝 and 퐶푔 are both measured from the tip of the nosecone).
Table 7 - Actual Stabilities
푪품 (inches) 푪풑 (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 퐶푝 (in red) and
퐶푔’s (in blue).
Figure 31 - Actual Cp and Cg locations
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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
46 | 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. 47 | P a g e
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
Percent Overall Weight Ballast Predicted Apogee Actual Apogee Difference (lb) (lb) (ft) (ft) (%) 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
48 | P a g e flight data was calculated, as well as error between the best fit line and the OpenRocket flight data.
|푡ℎ푒표푟푒푡𝑖푐푎푙−푒푥푝푒푟𝑖푚푒푛푡푎푙| % 퐸푟푟표푟 = × 100% (2) |푡ℎ푒표푟푒푡𝑖푐푎푙|
6000
5000
4000 Actual Altitude (ft)
3000
OpenRocket Altitude (ft) Altitude(ft) 2000
1000
0 0 20 40 60 80 100 120 140 160 Time (s)
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.
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6000 12%
5000 Regression Altitude (ft) 10% Actual Altitude (ft) 4000 Percent Error 8%
3000 6% Altitude(ft)
2000 4% PercentError(%)
1000 2%
0 0% 0 20 40 60 80 100 120 140 160 Time (s)
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.
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6000.00 12%
Regression Altitude (ft) 5000.00 10% Percent Error
4000.00 8%
3000.00 6% Altitude(ft)
2000.00 4% PercentError(%)
1000.00 2%
0.00 0% 0 20 40 60 80 100 120 140 Time (s)
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
51 | 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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6000 0.7
5000 0.6
0.5 4000 0.4 3000 0.3
Altitude(ft) Altitude (ft) 2000 Drag coefficient 0.2 Coefficient ofDrag
1000 0.1
0 0 0 5 10 15 20 Time (s)
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):
퐷 = 푚(푎 − 푔) (3)
Following the calculation of drag force, coefficient of drag can be calculated using the following equation:
퐷 퐶퐷 = 푣2 (4) 휌 퐴 2 푐
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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
54 | P a g e 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.
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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
58 | 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
59 | 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 Little to no reduction in overall safety of team members or 4-Negligible 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
60 | P a g e 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
Severity/ Post-Control Risk/Hazard Root Cause/Effect Mitigation and Control Likelihood Severity/Likelihood Open containers of epoxy during fabrication operations leading to inhalation of toxic fumes, Epoxy Fumes 4A Work in well ventilated spaces 4C accidental ingestion, or contact with skin leading to potential for irritation or rash Individuals handing epoxy must be wearing Proper PPE, such as gloves, pants, and close- Mishandling of epoxy during Epoxy toed shoes when handling epoxy to prevent application leading to skin 4A 4C Contacting Skin contact with the skin. In the event that epoxy irritation does come in contact with the skin, wash it off at the sink Mishandling of epoxy resulting in Handle the epoxy carefully during mixing or epoxy hardening on the working transport. In the event that any epoxy does Spill/Leak of area, potentially ruining lab 4C spill, wipe up the excess with a cloth and 4D Epoxy equipment or various parts of the dispose of it properly, and clean the dirtied launch vehicle area. Never leave epoxy unattended. Monitor the Epoxy Burning Mishandling of epoxy leading to heat of the epoxy as you mix it. If epoxy does Through potential damage to user, lab, or 2E 2E get excessively hot, remove sample from lab Container equipment and let it cool before disposing of it properly
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Table 13 - Personnel Hazard Analysis - Launch Operations/Post-Launch Inspection
Severity/ Post-Control Risk/Hazard Root Cause/Effect Mitigation and Control Likelihood Severity/Likelihood
Wear proper PPE, such as safety glasses Particles flying through the air during launch and fabrication. In the event Debris In Team during fabrication operations 2C that debris does enter the eye, the eyewash 3D Member's Eye leading to potential scrape or cut station will be used to cleanse the eye of the to user's eyes debris.
Team member will be required to wear proper Improper sanding or fabrication of Sharp Edges on PPE, such as gloves, close-toed shoes, and fins and nose cone resulting in Fins and 4D pants during testing operations and inspection 4E potential splinters or cuts to team Nosecone procedures to prevent direct contact with members fragments of the rocket
Improper fabrication operations or Team member will be required to wear proper Cracks or faulty components putting team PPE, such as gloves, during testing setup and Chipping in members at risk for splinters or 3C 4E inspection procedures to prevent direct Body Tube cuts when coming in contact with contact with fragments of the rocket rocket
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Materials Failure of component durability or Team members will wait in designated safe Experience subsystem resulting in a range of launch zone until rocket is deemed safe for Explosive possible injuries to team members 1E retrieval by RSO. Safety officer will retrieve 2E Breaking When from minor to severe depending on rocket, wearing proper PPE and keep face Opening for the intensity of the explosion directly out of line of launch vehicle. Inspection Oversight or ignorance when Direct Contact approaching hot materials for Proper PPE, such as gloves or aprons, must be With Hot 2D 4D handing, yielding to varying worn at all time when handling hot objects Material degree of burn to team members
Improper storage of flammable All flammable objects will be kept in proper components or inappropriate locations away from sparks and open flames. Materials fabrication operations/tool usage In the event of a small fire, a fire extinguisher 1D 2E Catching Fire leading to potential severe injury will be used to put out the fire. In the event of or burns to team members, a large fire, the team will evacuate the equipment, or work space building and the fire department will be called
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High exposure to black powder Black powder is stored in portable fireproof when handling and preparing case to keep away from fire and high Black Powder samples of this toxic gas can result 2D temperatures. When handling substance 3D Fumes in coughing, dizziness, and recovery subsection lead will measure fainting samples in well ventilated areas
Per the motor preparation checklist, the motor Rocket will be transported from an offsite location to Propellant Improper transportation and the launch location in a protective, waterproof Comes In configuration of motor subsystem 2C casing. Upon installation, the propulsion team 2D Contact With leading to irritation and burns lead will prepare the motor according to Skin manufacturer specification while wearing proper clothing, shoes and PPE
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Table 14 - Personnel Hazard Analysis - Testing
Severity/ Post-Control Risk/Hazard Root Cause/Effect Mitigation and Control Likelihood Severity/Likelihood Per malfunctioning electronics Unsafe working conditions or lack troubleshooting checklist, electronics of care when handling electronics subsection lead will inspect faulty instrument Electrical Shock leadings to electrocution resulting 3D for improper connection. Care has been take 3E in burns, significant injury, or to ensure nothing with exposed or fraying death wiring is being used in fabrication. Electronics will be stored in dry, secured area Improper handling of shop tools or Only authorized individuals have run tests. Inexperienced machining operations leading to 3C Multiple team members are present during 3D Test Personnel personal injury or destruction of testing to report and issue if one should occur equipment Team member have been required to wear Fractured Failure of various components proper PPE during testing setup and Particles During leading to potential splinters or 3B 4E inspection to prevent direct contact with Testing cuts to team members fragments of the rocket
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Table 15 - Personnel Hazard Analysis - Fabrication
Severity/ Post-Control Risk/Hazard Root Cause/Effect Mitigation and Control Likelihood Severity/Likelihood
Proper clothing, shoes, and PPE must be worn Allergic Handling of materials team at all time when handling materials. Allergies Reaction to member is allergic to resulting in of all team members are kept on file, and 2E 2E Building an allergic reaction in the form of members allergic to a specific material will Material skin irritation, rash, or swelling not work with that material while it is fabricated.
All team members have been trained on how Improper handling of shop tools or Improper Heavy to properly use shop equipment and have machining operations leading to Machinery 2C passed written and practical tests regarding 2D personal injury or destruction of Usage proper handling and maintenance of shop equipment equipment
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All team members have been trained on how to properly use handheld tools. During Improper Bruises, cuts or scrapes from fabrication operations, team members have a Handheld Tool mishandling of basic handheld 3C 3D spotter to ensure proper safety procedures are Usage shop tools such as hammer or saw followed and to monitor surroundings during fabrication operation
Tools are stored in proper locations to keep Tool storage in improper location team members and work area clean and safe, following fabrication operations Improper Tool and prolong life of tool. Periodic checks will leading, or usage by unauthorized 3C 3D Storage be conducted by safety officer to ensure all individuals leading to damage to materials are returned following construction equipment or environment. and placed in their proper locations
Cuts leading to injury as a result of During fabrication operations, team members Improper Use unsafe precision cutting operations have at least one spotter to ensure proper of Craft/Exacto 2D 2E on fins or other pieces of the cutting procedures are being followed by Knife rocket body cutting away from body.
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Proper clothing, shoes, and hair styles will be Lack of education on proper required in the lab to ensure safety for all Improper Work clothing or inspection of shop 4D team members. Safety officer will conduct 4E Attire workers leading to potential periodic checks of fabrication work attire and damage to body or attire PPE.
Cords or other materials lying on Cords will be plugged in closest to the area in the floor could cause team Tripping which their machine is being used. The work members to trip, thus resulting in 4B 4D Hazards area will be kept tidy in order to prevent cuts, scrapes, bruises, or broken debris from accumulating on the floor bones
Lack of awareness to surroundings Ensure all needed materials are close by Overreaching leading to potential falls, cuts, or 4B before beginning fabrication to avoid 4C scrapes overextension. Keep proper footing/balance.
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Table 16 - Personnel Hazard Analysis - Education Engagement Outreach Events
Severity/ Post-Control Risk/Hazard Root Cause/Effect Mitigation and Control Likelihood Severity/Likelihood Seat belts are worn at all times by all Not following driving rules and members inside the vehicle. All individuals in regulations resulting in a range of the vehicle will also sign a waiver releasing potential injuries to team members Car Accident 1E the team of liability in the event of an 2E in the car, or others from minor to accident. The driver must follow all federal severe, and potential property driving laws including have a valid license damage and insurance Lack of oversight of individuals Age appropriate tools will be given during managing event or disobedient Child Using educational outreach events. Strict participants could lead to child Tools 2D supervision will be used to monitor all 3D experiencing a range of injuries Inappropriately activates to ensure all children are safe and depending of the tool being used at know what they are supposed to do. the event Lack of oversight of individuals All children will be closely monitored in Child Not managing event or disobedient order to ensure they are doing what they are Following participants could lead to child 3D 4D supposed to. If they continue to be defiant, Instructions could experience a range of they will be removed from the activity. injuries depending operation/event
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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
Severity/ Post-Control Risk/Hazard Root Cause/Effect Mitigation and Control Verification Plan Likelihood Severity/Likelihood
Faulty components or
inappropriate fabrication Care has been taken to ensure all In accordance with the final
operations lead to failure of parts have been fabricated assembly checklist, each
Cracking or rocket upon launch or testing according to specification. Parts subsection lead will inspect their
Chipping of operation. Potential to splinter 2D will be stored in appropriate section of the rocket for any 3D
Fabricated Parts and cause significant damage to containers and holders within compromises in structural
other sections of the rocket or locked room to prevent integrity as a result of fabrication
lead to failure of subsequent accidental damage. operations
components
Fabrications not completed Only trained individuals are In accordance with the final according to specifications allowed to operate any assembly checklist, each Lack of Precision leading to potential inability to machinery during the fabrication subsection lead will inspect their When Fabricating assemble components of rocket 2D and construction process. Other 3D section of the rocket to ensure all Parts properly and have secure team members will verify work parts are fabricated to specified attachment, resulting in failure to ensure it meets standards set dimensions of rocket upon launch or testing. for in the design
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In accordance with the final All components of the rocket Fabrications not completed assembly checklist, each have been measure after according to specifications subsection lead will inspect their fabrication in order to ensure resulting in inability to assemble section of the rocket to ensure all they meet the dimensions Gaps Between components of rocket properly parts are fabricated to specified 2C specified in the design. The pre- 2D Connecting Pieces and have secure attachment dimensions. In the event that launch safety checklist will be potentially leading to failure of there are gaps between adjoining used to ensure team members rocket upon launch or testing sections, the troubleshooting visually inspect connections of operation. checklist will be followed to components prior to launch remedy the issue
In accordance with the secure Epoxy has been mixed in Lack of attention to security of attachment inspection within the accordance with instructions in connection causing inability of final assembly checklist, each order to ensure a good adhesive Insufficient Epoxy components to hold together 3D subsection lead will inspect their 3E mixture. Components will be leading to separation and section of the rocket to ensure tested prior to launch to ensure a potential failure attachment between adjoining secure, water-tight seal subsections
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Only trained individuals are
Lack of knowledge by team allowed to operate any
members during fabrication machinery during the fabrication All team members will pass a Wrong Equipment operations leading to potential and construction process. If a practical and written test on Usage for damage to component(s) of component is corrupted, 4D proper usage of shop equipment 4E Fabrication rocket and individual harm. Also fabrication will be done to before they are allowed to use Operation potential to generate flawed salvage as much of the material equipment. component that is not suitable as possible without
for usage compromising safety of the
launch vehicle and operations
Improper storage or handling of Team members will operate in a
materials causing potential safe manner to prevent the Team members will be trained damage to component(s) of start/spread of fire. In the event on how to properly use fire rocket and individual harm of a small fire, it will be extinguisher in the event of a Materials Catch Fire resulting in minor to major loss 1D extinguished using the fire 2E small fire. Safety officer will of equipment, workspace, or extinguisher in the energy periodically test fire extinguisher components, or compromise of systems lab. For larger fires, 911 to ensure it is fully functional. structural integrity of launch will be called and the team will
vehicle retreat to a safe distance.
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Lack of knowledge as to where The parts checklists will be used to put supplies when fabrication The energy systems lab is to sign tools and materials in and operation is completed leading cleaned after each work period. out. Additionally, the safety to potential damage to Checklists have been created to Improper Storage of officer will periodically check materials/equipment resulting in 4A ensure that all materials and 4B Materials/Equipment the supply cabinets to ensure all compromise of the structural equipment being used are tools are returned and in their integrity of various components, returned to their proper locations proper locations following or inability to use equipment for before everyone can leave. fabrication operations further fabrication operations
Parts checklists will be used to
check out tubes of epoxy so that Oversight of connection security the safety officer has all supplies during inspection process or Epoxy is stored in the in the lab accounted for. Furthermore, each Degradation of improper storage leading to lack at room temperature according 3E subsection of the rocket will be 3E Epoxy of adhesion between parts to specification listed by the inspected via the final assembly resulting in separation and manufacturer. checklist to ensure proper potential failure connection and adhesion
between adjoining sections.
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Table 18 - Failure Modes and Effects Analysis - Payload
Severity/ Post-Control Risk/Hazard Root Cause/Effect Proposed Mitigation Verification Plan Likelihood Severity/Likelihood
Failure to properly prepare ignition system according to Black powder is stored in a dry checklist leading to a range of area at room temperature in failure modes from minor accordance with manufacturer See ejection testing summary Premature Ignition damage to the payload system or 2E specifications. Testing has been and results in project plan 3E Charge components to catastrophic done to ensure premature section of FRR failure due to premature ignition does not lead to separation and parachute recovery failure deployment
Testing and research has been completed in order to ensure the Faulty motor or improper proper motors for each size of Motor has been tested via full- storage leading to inability for launch vehicle created is being scale launch operations. For Failure of Motor rocket to ascend off launch pad. 3D 3E used. Rocket motors will be kept further detail see full scale Potential damage to payload or in a dry area at room testing section of FRR other components upon misfire temperature in accordance with manufacturer specifications.
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Improper storage of black Black powder is stored in a dry powder sample or compromised area at room temperature in sample resulting in an inability accordance with manufacturer See ejection testing summary Failure of Black for launch vehicle to separate 1D specifications. Testing has been and results in project plan 1E Powered Charge leading to potential catastrophic done to ensure proper amounts section of FRR damage to payload and rocket of black powder is used for failure of recovery system ignition.
Ejection test was completed in order to determine the proper Inability to input proper amount amount of black powder to be of black powder into launch Black powder is stored in a dry Deployment of used to pressurize the launch vehicle as determined by testing area at room temperature in Black Powder vehicle and deploy the resulting in over pressurized accordance with manufacturer Change Resulting in 1E parachutes. Impact tests were 2E capsule causing minor to specifications. Testing is done to Damage to Payload also completed on the payload catastrophic damage to payload, ensure proper amounts of black Holding Container container to determine how holding container, or spring- powder is used for ignition. much force is felt by the fragile damper system material while within the dampening system
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Failure to properly account for Tensile and impact tests were forces experienced by launch Testing has been conducted in completed on the payload vehicle during flight or improper order to minimize forces on container in order to measure it's Bending/Breaking of assembly resulting in payload and ensure fragile tensile and ability to dampen Spring-Damper 3A 3D compromise in the structural payloads of all types will be kept direct impact. For further detail, System integrity of the spring-damper safe and secure during launch see MTS testing summary and system and damage to the fragile and recovery operations results in project plan section of payload FRR
Failure to properly account for An inspection has been In accordance with the final forces experienced by launch completed in accordance with assembly checklist, the payload vehicle during flight or improper the pre-launch safety checklist in container will be inspected for Crack in Payload inspection prior to launch 3D order to ensure the structural cracking or any other structural 3E Holding Container resulting in compromise in the integrity of the payload systems imperfections that could have structural integrity of the spring- is not in any way compromised been acquired during fabrication damper system and damage to prior to launch operations or transport prior to launch the fragile payload
Dampening material is unable to Testing was completed in which absorb impact and restrict the movement of the payload Inability to Keep movement leading to potential within its holding container is See MTS testing summary and Payload Static minor to catastrophic damage to measured in order to ensure it 2D results in project plan section of 3D within Holding fragile payload and damage to does not experience collision FRR Container other components of the launch with surrounding walls or vehicle. Potential failure to meet anything else that could cause mission objective fracture or damage.
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Tests were conducted in order to Inability to slow speed of launch validate that the materials used Payload Damage vehicle during decent leading to for the payload security See MTS testing section of FRR Upon Impact With damage to fragile payload 2E container can withstand the for results on impact testing of 2E Ground Upon resulting in repair, replacement, forces experienced by the fragile payload container Decent or failure to meet mission material without damage to its objective structural integrity
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Table 19 - Failure Modes and Effects Analysis - Payload Integration
Severity/ Post-Control Risk/Hazard Root Cause/Effect Proposed Mitigation Verification Plan Likelihood Severity/Likelihood
In accordance with the final Throughout the fabrication assembly checklist, each Fabrication operations not process all complete parts have subsection lead will inspect their completed according to been measured and verified by section of the rocket to ensure all Lack of Space in specifications leading to inability the team lead in order to ensure parts are fabricated to specified Body Tube for to properly assembly launch 3C 3E they meet the proper dimensions dimensions. In the event that Payload Container vehicle resulting in removal of laid out in the design. This will there are gaps between adjoining payload container and failure to allow for proper fit and connect sections, the troubleshooting meet mission objective within the launch vehicle checklist will be followed to remedy the issue
Failure to properly inspect payload subsection system prior Prior to launch the pre-launch to launch causing potential In accordance with launch Payload Container checklist will be used to verify damage to the payload or its procedures checklist, payload Not Properly that payload is mounted housing container. Could 2D will be reviewed for flight 3D Mounted in Body correctly in place and all compromise the structural readiness and proper mounting Tube connections are secure to ensure integrity of various components prior to launch operations safe launch operations or lead to failure of other operations
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In accordance with the secure attachment inspection within the Failure to properly inspect final assembly checklist, the Prior to launch the pre-launch connection between adjoining connection of adjoining Weak Attachment checklist will be used to verify subsections leading to possible subsections will be checked by Between Payload that payload is mounted cracking or separation between 2D the safety officer to ensure 3D Container and correctly in place and all the payload and recovery proper connection. In the event Recovery System connections are secure to ensure systems and inability to return that there are gaps between safe launch operations fragile material adjoining sections, the troubleshooting checklist will be followed to remedy the issue
Failure to fabricate payload All fully fabricated parts have container according to been measure and compared to Payload container has been Inability to Fit specifications given by NASA design requirements in order to design in accordance with the Given Payload Into resulting in potential inability to 3E ensure they meet the proper 3E envelope of fragile material Container meet mission objective of safely dimensions, thus ensuring that provided by NASA launching and returning a fragile the fragile payload fits within the material on our launch vehicle envelope of the container
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Improper material used to Inability to Fill dampen forces or unreliable Testing has been done in order Payload Container See MTS testing section of FRR impact testing data causing to measure the force felt by the with Material that 2C for results on filler material's 3D damage to the payload that could payload during launch and Dampens Force Felt ability to dampen impact result in failure to meet mission landing operations by Payload objective Inability to Fill Improper material used to Testing has been done in order Payload Container restrict movement or unreliable to measure the movement of the See MTS testing section of FRR with Material that impact testing data resulting in payload within the container in 3C for results on filler material's 4D Restricts Payload damage to the payload that could order to ensure it will not be ability to restrict movement Movement During result in failure to meet mission damaged as a result of striking Flight objective interior walls
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Table 20 - Failure Modes and Effects Analysis - Recovery System
Severity/ Post-Control Risk/Hazard Root Cause/Effect Proposed Mitigation Verification Plan Likelihood Severity/Likelihood
Testing have been done in order to verify the packing method and Various parachute packing Improper packing method give the student practice with methods have been researched leading to failure in parachute to packing the parachute into the and tested in order to determine Parachute is Not deploy properly resulting in body tube. On launch day, the an optimal method. Tests have 1E 3E Packed Properly launch vehicle experiencing pre-launch checklist will be been conducted with various more force than planned upon followed to ensure proper packing styles. For further detail, when landing packing of the parachute in see parachute deployment force accordance with standard testing subsection of FRR practices.
Various tests have been Failure to properly inspect completed in order to verify the parachute prior to launch or In accordance with the recovery strength of the parachute. The faulty parachute resulting in preparation checklist, all parachute will be inspected prior launch vehicle descending at a parachutes will be inspected Tear in Parachute 2D to launch using the pre-launch 3E faster rate than planned in an prior to launch for tears, snags, checklist in order to verify it uncontrolled manner causing or any other imperfection that does not have any tears, pulls, potential damage to components could result in recovery failure rips, or other imperfections that or total loss could result in failure
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Tests have been run in order to Failure to properly inspect shock verify the strength of the shock cord prior to launch or faulty In accordance with the recovery cords. The cords will be component resulting in launch preparation checklist, all shock inspected prior to launch using vehicle descending at a faster cords will be inspected prior to Tear in Shock Cord 2E the pre-launch checklist in order 3E rate than planned in an launch for tears, snags, or any to verify it does not have any uncontrolled manner potentially other imperfection that could tears, pulls, rips, or other damaging components or result in recovery failure imperfections that could result in resulting in a total loss failure
Testing has been done in order Incorrect shock cord for force to verify the strength of the experienced during deployment connection between the shock See parachute deployment force Shock Cord Cannot causing separation of rocket into cords and the main rocket body testing subsection of FRR for Withstand Force of multiple pieces, some of which 1D tube. The connection between details regarding shock cord 2D Parachute will not be attached to the these two will also be inspected strength and durability for Deployment parachute, causing damage and prior to launch with the pre- various black powder charges potential harm to by standards launch checklist in order to ensure a secure attachment
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Failure to properly pack parachute or use correct amount of black powder for See parachute deployment Tests have been conducted in pressurization leading to testing and full-scale testing order to verify the precise uncontrollable decent until the subsections of FRR for details Drogue Parachute amount of black powder that will opening of the main parachute 1E regarding proper quantity of 3E Deployment Failure need to be used to pressurize the resulting in the launch vehicle black powder for launch vehicle parachute and allow for proper landing with a greater impact, pressurization and drogue deployment. thus causing damage to parachute deployment components or endangering spectators
Failure to properly pack parachute or use correct amount See parachute deployment Various tests have been of black powder for testing and full-scale testing completed in order to verify the pressurization leading to decent subsections of FRR for details Main Parachute precise amount of black powder at a quicker rate than expected 1E regarding proper quantity of 3E Deployment Failure that will need to be used to and potential drift of the launch black powder for launch vehicle pressurize the parachute and vehicle off course resulting in pressurization and main allow for proper deployment. damage to the components or parachute deployment endangerment of spectators.
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Launch of rocket in excessively The rocket will only be launch in windy conditions resulting in an proper conditions, therefore inability for parachutes to open UE SLI team in conjunction with minimizing the chance for wind at proper altitudes or launch at NASA and local rocket clubs Wind Blows Rocket gusts to blow the rocket off excessive wind speeds leading to 3D will monitor wind speeds and 4E Off Course course. In the event that this potential for rocket to become make launch related decisions does occur, the launch vehicle lost or components becoming accordingly will be retrieved using the GPS damaged if landing occurs tracking system in the altimeter outside of the space given. Testing has been done in order to verify the precise amount of See altimeter testing and full- Incorrect packing of parachute, black powder that will need to scale testing subsections of FRR or faulty electronics leading to be used to pressurize the for details regarding proper Parachute Deploys potential for uncontrollable parachute and allow for proper 2D quantity of black powder for 3D at Incorrect Time decent, damage to components deployment at the correct time. launch vehicle pressurization or compromises to structural The recovery system will also be and parachute deployment at integrity of the launch vehicle tested prior to launch in proper times accordance with the pre-launch checklist. Interference from Operating on the same, or close Testing the recovery system with Verify that all electronics are the Scoring to the same frequency resulting the scoring altimeter on and 1D working according to the Pre- 1E Altimeter Causes in a failure in the recovery nearby to determine if there is Launch checklist. System Failure system. any interference with the system.
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Table 21 - Failure Modes and Effects Analysis - Testing
Severity/ Post-Control Risk/Hazard Root Cause/Effect Proposed Mitigation Verification Plan Likelihood Severity/Likelihood Lack of knowledge or All tests have been done by experience by test personnel supervisors of each group. Subsection team leads will be Inexperienced resulting in damage to lab Multiple team members will be presents thought the duration of Student Running equipment, facilities, 3D present during testing in order to 4D their area's testing to ensure Tests components of the launch ensure proper protocols are proper usage of lab equipment. vehicle, or other team members followed and safety precautions or students are taken
All tests have been done by Lack of knowledge or Aerodynamics subsection team supervisors of each group. experience by test personnel leads will be presents thought Wind Tunnel Multiple team members will be resulting in damage to the testing the duration of wind tunnel Operation at 4D present during testing in order to 4E apparatus or components of the testing to ensure proper usage of Excessive Speeds ensure proper protocols are rocket being tested in the wind lab equipment in accordance followed and safety precautions tunnel with manufacturer specifications are taken
Failure to properly clean and Proper PPE and eye protection inspect wind tunnel prior to must be worn at all times in the Safety officer will monitor testing resulting in potential lab. In the event that debris does testing to ensure proper PPE, Debris in the Wind damage to testing apparatus, 4B fly out of the wind tunnel during such as gloves, ear and eye 4D Tunnel components of the rocket or testing, multiple students will be protection are being worn during harm of individuals running the present to assist in the clean-up testing operations. test of the debris.
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Wind tunnel testing will be Aerodynamics subsection team Excessive testing beyond scheduled in advance with leads will be presents thought apparatus capabilities causing breaks in between tests to allow Overuse of Wind the duration of wind tunnel damage to the testing apparatus 3D the engine to properly cool. The 3E Tunnel testing to ensure proper usage of that could result in inability to tunnel will also be inspected lab equipment in accordance conduct future tests before testing to ensure proper with manufacturer specifications conditions.
Improper storage of black powder leading to no separation Secondary charges can be used See ejection testing subsection Black Powder Fails or deployment of parachute, thus in order to ensure that if one 2D of FRR regarding proper black 3D To Ignite creating a potential for change fails another can engage powder handing catastrophic damage to launch to deploy the parachutes. vehicle or injury to spectators
Failure to properly measure correct amount of black powder Manufacturer specification are See ejection testing subsection Excess of Black for sample resulting in full followed in order to determine of FRR detailing the amount of Powder Used in separation of rocket leading to 2D how much black powder is need 2E black powder to be used for Testing damage to various components to pressurize the rocket based on complete and optimal separation and potential failure of other its weight systems and debris
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Improper connection inspection prior to launch leading to Prior to testing the pre-launch In accordance with launch potential damage to the payload checklist were used to verify that procedures checklist, the Failure to Properly or its housing container. Could payload is mounted correctly in 3C payload will be inspected prior 3D Secure Payload compromise the structural place and all connections are to launch for secure attachment integrity of various components secure to ensure safe launch and flight readiness or lead to failure of other operations operations Multiple tests have been Testing have been done in order conducted in order to verify the to verify the packing method and Incorrect packing procedure or packing method and give the give the student practice with method used resulting in failure student practice with packing the packing the parachute into the Parachute is Not in parachute to deploy at proper parachute into the body tube. On body tube. On launch day, the Packed Properly for 1C 3D altitude resulting in launch launch day, the pre-launch pre-launch checklist will be Testing vehicle experiencing more force checklist will be followed to followed to ensure proper than planned upon landing ensure proper packing of the packing of the parachute in parachute in accordance with accordance with standard standard practices. practices.
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Table 22 - Failure Modes and Effects Analysis - Launch Support Equipment
Severity/ Post-Control Risk/Hazard Root Cause/Effect Proposed Mitigation Verification Plan Likelihood Severity/Likelihood
Faulty component(s) or failure to In accordance with set up on account for various changes to Prior to launch operations, the launch pad checklist, the launch rocket made between launches guide rails will be inspected by pad and guide rail will be Instability in Guide can cause launch vehicle to the lead safety officer in inspected for structural flaws or 2C 3D Rail deviate from its deal path accordance with the pre-launch bowing that could lead to potentially leading to checklist in order to ensure safe instability in launch and endangerment of spectators or operations deviation of the rocket from its damage to components ideal flight path
Lack of care when handing The launch vehicle will be In accordance with final launch vehicle or components transported in its specially made assembly checklist, prior to resulting in potential damage to container which will provide launch, all subsection will be Improper Transport launch vehicle and/or 2C support for all fragile areas of inspected by their team lead for 3D of Launch Vehicle compromise of structural the rocket, while protecting it cracks, chipping, or other integrity of individual from slipping, vibrations or other structural flaws that could have components potentially damaging impacts. been acquired during transport
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The launch vehicle is stored in a specific container in a locked In accordance with the parts Lack of care and safety when room during fabrication and in checklist, the safety officer will storing launch vehicle or between tests. Only team leads periodically inspect storage Improper Storage of components leading to damage and safety officers have access cabinets as well as launch 2D 3D Launch Vehicle to launch vehicle and potential to the room to prevent the rocket vehicle holders to ensure all compromise of structural from being mishandled. The supplies have been returned integrity of components room is kept at room following use, and are being temperature to not adversely stored in their proper place affect any components
The rocket motors will be transported in a fireproof case In according with motor Handling of motor not in that will prevent moisture for preparation checklist, the rocket accordance with specifications Improper Transport getting into the motor. The case motor will be stored off site and leading to potential damage to 3C 3E of Rocket Motor will also protect the motors will be transported to the launch payload or other essential against slipping, vibrations, and location in a protective, components upon launch other potentially damaging waterproof case impacts.
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The rocket motors are stored in a fireproof case. This case is kept Potential damage to payload or In according with motor in a locked cabinet in order to other essential components upon preparation checklist, the rocket prevent from other team Improper Storage of launch due to improper motor will be stored off site and 2D members handing the motors. 3E Rocket Motor placement of motor in will be transported to the launch This cabinet will remain at room potentially compromising location in a protective, temperature and dry in order to locations waterproof case not allow heat or moisture adversely affect the motor.
In accordance with motor Handling of launch vehicle not preparation checklist, the leader Only trained and essential team in accordance with guidelines of the propulsion subsection will members will handle the rocket Improper Handling listed on pre-launch checklist retrieve the rocket motor from its during launch operations. Pre- of Rocket on Launch leading to endangerment of 2D protective, waterproof casing, 2E launch safety checklists will be Pad spectators, and minor to and will ready the motor for used in order to ensure everting catastrophic failure of the rocket ignition after all subsequent is safe for launch and its subsystems subsection inspections have been completed
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In accordance with the set up on Faulty component or failure to launch pad checklist, prior to Prior to launch operations, the properly inspect launch pad prior launch operations, the launch launch pad will be inspected by to flight leading to potential for pad will be retrieved from Instability of Launch the lead safety officer in rocket to deviate from ideal 2B NASA will be inspected by the 2C Pad accordance with the pre-launch flight path endangering lead safety officer in conjunction checklist in order to ensure safe spectators and causing failure of with the RSO for any structural operations components or drift flaws that could lead to instability in launch operations
Prior to launch operations, the In accordance with ignition Improper handling or storage of ignitor clips will be inspected by checklist, ignitor clips will be component causing rocket to be the lead safety officer in Faulty Ignitor Clips 3C inspected prior to attachment to 3D unable to ascend off of launch accordance with the pre-launch the launch vehicle by propulsion pad checklist in order to ensure safe subsection lead. operations and successful launch
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Table 23 - Failure Modes and Effects Analysis - Launch Operations
Severity/ Post-Control Risk/Hazard Root Cause/Effect Proposed Mitigation Verification Plan Likelihood Severity/Likelihood
Improper storage of launch Prior to launch, the body tube
vehicle or transportation leading will be thoroughly inspected for In accordance with the launch
to compromise in the structural cracking, splintering, or procedures checklist, the main Cracking in Main integrity of the rocket leading to 2D fatiguing in according with the body tube will be inspected by 3D Body Tube potential damage to other procedures listed in the pre- the safety officer prior to launch
components or failure of other launch checklist in order to for any structural imperfections
subsystems ensure safe launch operations
All components of the rocket All components of the rocket Failure to fabricate subsections were measured after fabrication have been measure after according to specifications in order to ensure they met the fabrication in order to ensure yielding an inability to assemble dimensions specified in the they meet the dimensions Gaps Between components of rocket properly 2C design. The pre-launch safety specified in the design. The pre- 3C Connecting Pieces with secure attachment checklist will be used to ensure launch safety checklist will be potentially leading to failure of team members visually inspect used to ensure team members rocket upon launch or testing connections of components prior visually inspect connections of operation. to launch components prior to launch
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Test have been run in open areas
to ensure no overhanging trees,
roofs, or other items could
Failure for parachutes to deploy impede the flight operations.
at proper altitude resulting in Furthermore, subsequent tests See scale model testing Collision with damage to launch vehicle and have been completed in order to subsection of FRR report Object in Sky (Tree, 2C 2D compromise in structural determine the strength of the regarding impact testing of Bird, Etc.) integrity of impacted body tube and nosecone so that launch vehicle
components in the event the launch vehicle
does strike a bird it can
withstand impact and return
safely
Failure of team members to
account for ballast adjustments In accordance with post-flight
made prior to launch resulting in inspection, all team members Instability During Maintain safe distance from change of center of gravity and 1B will wait until rocket has landed 2C Flight launch pad leading to inability of the rocket in a safe location before leaving
to maintain its projected flight safe launch zone
path
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In accordance with launch Failure to properly inspect procedures checklist, all Altimeter or Other security of attachment between Verify all electronics work electronics will be tested for Electronics in electronics, or test functionality 3B properly before launch and are functionality prior to launch 3C Avionics Bay prior to launch causing potential firmly attached to the rocket operations. For further detail on Malfunction/Fall Off short circuiting or harm to electronics testing, see altimeter spectators below testing section of FRR
In accordance with launch
Parts not fabricated according to procedures checklist, the ability Run multiple tests to ensure specifications resulting in of adjoining sections to separate Coupler Excessively proper amounts of black powder potential failure of parachute to 2D will be tested. For further detail 3D Tight is used to allow rocket to deploy leading to damage to on proper ejection charges for separate rocket separation, see ejection testing
subsection of FRR
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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
Root Cause/ Severity/ Mitigation and Post-Control Risk/Hazard Effect Likelihood Control Severity/Likelihood Vehicle Effects on Environment
Fumes released during construction Work in well resulting in ventilated spaces Epoxy Fumes hazardous working 4A 4D and dispose of conditions created waste properly for team members as a result of toxic air
Failure to follow proper disposal Let epoxy fully Epoxy Not protocols leading cure before Disposed of to potential fire 4C disposal in order 4E Properly hazard and damage to prevent fire to lab or hazard equipment
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Fabrication operations Wear mask when producing small sanding to avoid dust particles from inhaling dust sanding or particles and try Dust Particles machining 4A to contain dust 4D operations are when sanding released into the opposed to freely environment which releasing it into can result in surrounding air. breathing problems Failure to properly secure motor, therefore, upon Place flame ignition, when resistant material motor reaches high beneath the temperatures and Rocket Motor launch pad to hot exhaust is 2D 3D Ignition avoid burning the released, the motor immediate could become surroundings or displaced or burn starting a fire the areas where the rocket is launched or lands
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Ensure fully Component failure functioning leading to parachutes before fragments of the launch via pre- rocket breaking off launch recovery during flight or preparation upon landing checklist and impact and check to make Debris from becoming sure all 3D 3E Rocket irretrievable, components of leading to minor the rocket and environmental payload are harm due to accounted for inability to upon return in decompose and accordance with toxicity of post-flight component inspection checklist
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Environmental Effects on Vehicle Improper storage or launching in unfit conditions can lead to water exposure, which Avoid launching can cause rocket in wet malfunctioning of conditions and electronics within store rocket in Water 2E 3E the avionics bay, proper stand in a or damage to the dry area for body of the rocket, storage and which will be transport constructed out of Blue Tube that is not 100% water resistant
Launch into excessive wind speeds leading to deviation from Avoid launching launch vehicle's rocket on days of Wind ideal flight path 3C high speed winds 3E thus leading to or unpredictable, damage to the strong wind gusts rocket and potential harm to spectators
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Improper storage of components resulting in potential corrosion Store rocket in a and weakening of dry area to avoid various materials moisture entering Humidity/Moisture 3D 4E used to construct the rocket over the rocket. This time via humid moisture can also air negatively impact on-board electronics Launch during times of low cloud coverage resulting Avoid launching in inability to track rocket on days Visibility the rocket thus 4C with low cloud 4E leading to debris coverage and not being retrieved poor visibility and damaging the environment
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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
Root Cause/ Severity/ Mitigation and Post-Control Risk/Hazard Effect Likelihood Control Severity/Likelihood
Being a first year team with a small The team has work budget could lead with faculty members to a lack of quality as well as local design or Limited rocketry club members fabrication 2C 2C Resources in order to gain a better material and to understanding of failure to meet rocketry and develop a mission objective functional rocket. or overall poor performance
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The team and its adult Being a first year educators have applied team with minimal for and been given established grants in order to fund funding, the team parts of the project. could be forced to Tight or Additionally, the team use parts that are 3A 3B Minimal Budget has held fundraisers to not optimal, or be provide the team with a unable to replace flexible budget beyond parts of the rocket the normal amount of that are broken money allotted to the during testing project by the school
Inability to manage project on Team members have a weekly basis and will continue to fill could potentially out weekly time cards lead to major Mismanagement and log their hours in delays resulting in 1E 2E of Time the task breakdown in the quality of work order to ensure lacking, or the everyone remains on rocket not being schedule completed by competition
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Inability to budget There has been and will time properly continue to be constant could leave the communication Underestimation project running amongst all team of Scope of behind schedule 2E members and with 3E Work and various facets NASA project leads to of the rocket not ensure the scope of being completed in work is clear and a quality manner everyone stays on task Failure to meet The team has designed proper FAA and and downselected with NASA safety safety as the foremost regulations could priority, and will lead to team to be Increase in clearly identify all forced to add Safety 2D safety measures before 2E material to the Regulations all operations so that rocket in order to additional, last-minute increase safety, safety measures do not which will result have to be taken that in an increase in will inflate the budget. expenses
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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
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Initial Part Quantity
Water -
Rags 3
Pocket Knife 1
Flat head screw Driver 1
Super Lube Synthetic 2
Grease
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 12
Isolators
5.36" Blue Tube (Cylinder 1
2)
Base Springs (#866) 5
Spacer (clear acrylic) 1
Recovery Bolts 3/8" x 1.25" 2
Length
Spring Fastening Bolts For 30
3/8" x 16 Bolt, 1/4" Height
Spring Fastening Washers 30
Bulkheads 2
Pins 0.1405 inch diameter 24
Aluminum Squares 1x1x0.1 12
inches
3D Printed Cylinder 1
(Cylinder 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 4
Ballast
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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
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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
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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
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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
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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
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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
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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
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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
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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
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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
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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
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Initial Check-Off Point
Ignitor and ignition system is set-up and ready for launch
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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
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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 All three altimeters precisely measured altitude. Altitude, GPS tracking, Altimeter Testing The GPS tracking and live and live feed feed worked properly. Complete. The epoxy failed, not the The force required for MTS Bulkhead Testing carbon fiber or aluminum. failure of the assembly. Complete. If separation is achieved, Separation was achieved Ejection Testing the amount of black powder 10 times for each body tube. needed. Complete. Parachute Force Force of the parachute In Progress. Deployment Testing deployment. Strain from a strain Wind Tunnel Testing Incomplete. gauge. Two Successful Flights. Scale Model Testing Full System Test Complete. Complete spring constant Payload Spring Testing Spring Constant Check matches. Complete.
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Test Data Taken Status and Results Three Successful Flights. Full Scale Testing Full System Test Complete. Modified Charpy Impact Payload impact testing Complete Test
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
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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 134 | 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
Number of Number of Signal Notes Tests Failures
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
Maximum Flight # 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 Width Model number Diameter (in) Length (in) Make/Model (in)
Vishay -TN-505-4 - Strain gage - - - Strain Gage
Strain Vishay 3800 Wide - 0.5 - indicator Range Strain Indicator
Scale Model 0.5
Air fan 144924 10
Wind Tunnel - - 135 -
- Test section 10 15 -
- Motor 1295L108A - - -
Differential Honeywell - pressure SSCSNBN010NDAA5 - - - transducer
Cantilever 6061 rectangle - 7.125 1 beam Aluminum beam
Conditioner m-prep conditioner A - - -
m-prep neutralizer Neutralizer - - - 5A
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Instrument Width Model number Diameter (in) Length (in) Make/Model (in)
320 and 400 grit Carbide paper - - - 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
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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.
Fd CD = U2 (1) ρ A 2 c
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3 Where Fd is the drag force(lbf), ρ is the air density ( lbm/ in ), 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.
ϵEwt2 F = (2) d 6L
Where ϵ is the strain (in/in), w is the width of the beam (in), t is the thickness of the beam
2 (in), E is the modulus of elasticity of the beam (lbf/in ), 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):
Eϵwt2 CD = U2 (3) 6Lρ A 2 c
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 = 5.202 ∆P (5) ( ) (in H2O) ft2
Therefore, the velocity is calculated using Equation (6).
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2 ∆P 2 U = √ (lbf⁄ft ) (6) 0.00226
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 Random (Precision)
(Bias) Uncertainty
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
100.00 90.00 80.00 70.00 60.00 50.00 40.00 Percent(%) Systematic (Bias)(%) 30.00 Random (Precision)(%) 20.00 10.00 0.00
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
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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.
푭 = −풌풙 Equation 1: Hooke’s Law The results of this test are summarized in Table 57.
Table 57 - Spring Constant Test Values
Spring Spring Mass Weight Displacement Displacement k (kg/m) k (lbf/in) (kg) (lbf) (m) (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
150 | 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
Fill Material Cotton Filling Shredded Paper Paper/Cotton mix Base Value Acceleration (counts/g) average 11444 2452 N/A 11643 of x,y and z directions. Percent reduction 1.7 78.9 N/A N/A from base
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
151 | P a g e 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% 퐹푆 ; Where FS= 32 g Equation 2
Nonlinearity (x,y,z)=±0.005 × 32 = ±0.16푔 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
2 2 Equation 6 Overall systematic error = B = BOIE=√푒 + 푒 퐿 푧
2 2 2 2 Equation 7 BOIE(x,y) =√푒 + 푒 = √(0.16푔) + (0.15푔) =±0.2193 푔 퐿 푧
2 2 2 2 Equation 8 BOIE(z) =√푒 + 푒 = √(0.16푔) + (0.25푔) =±0.2968 푔 퐿 푧
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
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
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
153 | 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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40000
30000
20000
10000
0
-10000
-20000 Acceleration Acceleration (counts/hz)
-30000
-40000 0 200 400 600 800 1000 1200 1400 1600 Time Step 4 Hz
X Y Z
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.
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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 Verification Status & Location Summarized Requirement Description of Verification Plan Number Method(s) The vehicle shall deliver the The rocket team will utilize Full scale test completed with science or engineering payload Test OpenRocket, RockSim, CFD, & 1.1 apogee of 5,291 feet. See “Full to an apogee altitude of 5,280 Analysis test flight data to achieve an Scale Flight” for more detail. feet above ground level (AGL). accurate prediction of altitude. The vehicle shall carry one The rocket will house a Atlus Three altimeters meeting commercially available, Metrum TeleMega altimeter in requirements were flown for barometric altimeter for 1.2 Inspection the nosecone to record the official full scale; all producing valid recording the official altitude altitude used in determining the altitudes. See “Validity used in determining the altitude altitude award winner. Assessment” for more detail. award winner. Recovery altimeters powered by All recovery electronics shall be Batteries & altimeter will be Energizer 9V Lithium Batteries. 1.3 powered by commercially Inspection purchased from online rocketry See “Line Item Budget in available batteries. sources. Appendix F”. The launch vehicle shall be The rocket is reusable in design Three test flights were designed to be recoverable and because the team is using a motor conducted on the February 18th. reusable. Reusable is defined as Test 1.4 that has refuels that can be No repairs were made to the being able to launch again on the Inspection reloaded into the motor under rocket, making it reusable. See same day without repairs or supervision. “Full Scale Flight”. modifications.
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NASA Requirements Handbook Verification Status & Location Summarized Requirement Description of Verification Plan Number Method(s) The launch vehicle has 3 The launch vehicle will have 3 The launch vehicle shall have a independent sections: the aft independent sections: the aft body 1.5 maximum of four (4) Inspection body tube, the bow body tube tube, the bow body tube and independent sections. and nosecone, and the coupler. nosecone, and the coupler. See “Vehicle Criteria”. The launch vehicle shall be Inspection The launch vehicle shall be a Only one L850W is used, as 1.6 limited to a single stage. Demonstration single stage. seen in “Vehicle Criteria”. The team was able to prepare The launch vehicle shall be The team will conduct multiple the rocket in 32 minutes on capable of being prepared for 1.7 Test tests on full-scale test day and February 18th. See “Full Scale flight at the launch site within 4 measure re-launch times. Flight” for more detail on hours. multiple launches that day. The launch vehicle shall be The launch vehicle design will All systems remained on during capable of remaining in launch- ensure all components have a life the full scale test for over 2 ready configuration at the pad for 1.8 Test of greater than 1 hour without hours while the rocket was a minimum of 1 hour without loss of functionality via a full- stuck in a tree. See “Full Scale losing the functionality of any scale launch pad test. Flight” for more detail. critical on-board component. The ignition system and igniters The launch vehicle shall be used during the full-scale test is capable of being launched by a Inspection The ignition system will use a 12- 1.9 12V. See the “Line Item standard 12-volt direct current Test volt direct current firing system. Budget” in the Appendix for firing system. exact ignitor specifications.
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NASA Requirements Handbook Verification Status & Location Summarized Requirement Description of Verification Plan Number Method(s) There will be no external circuity The ignition system comprised The launch vehicle shall require for the ignition system because it of only 1 ignitor that runs off of no external circuitry or special will be a ground based ignition 12V. This setup was successful 1.10 ground support equipment to Inspection system being placed underneath during the full scale test. See initiate launch (other than what is the rocket before launch with 300 the “Line Item Budget” in the provided by Range Services). ft of cord between the igniter and Appendix for exact ignitor the controller. specifications. The launch vehicle shall use a commercially available solid The team has purchased and motor propulsion system using The motor being used is a solid flown on an Aerotech L850W, ammonium perchlorate Inspection 1.11 fuel motor from AeroTech. The see the “Line Item Budget” for composite propellant (APCP) motor is the L850W. further information on the which is approved and certified motor. by the National Association of Rocketry (NAR). As of final design, no pressure Pressure vessels on the vehicle 1.12 Inspection No pressure vessels will be used. vessels are used. See “Design shall be approved by the RSO. and Construction of Vehicle”. The total impulse provided by a The motor will produce an The motor details can be found University launch vehicle shall impulse of 3695 N-s which is via Aerotech’s website. See 1.13 Inspection not exceed 5,120 Newton- below the specified total impulse “Line Item Budget” for specific seconds (L-class). that is allowed. motor look-up information. The launch vehicle shall have a Test Using OpenRocket, Rocksim, and The flight configuration for 1.14 minimum static stability margin Test Data – determine rail exit competition has an actual flight Analysis of 2.0 at the point of rail exit. velocity and then stability. stability of 2.70.
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NASA Requirements Handbook Verification Status & Location Summarized Requirement Description of Verification Plan Number Method(s) The rocket team will utilize OpenRocket, RockSim, CFD, & The Full-Scale flight was a The launch vehicle shall test flight data to achieve an success. See Section “Flight 1.15 accelerate to a minimum velocity Analysis accurate prediction of minimum Simulations and Altitude of 52 fps at rail exit. velocity at rail exit. The current Predictions” value is 66.9 fps. All teams shall successfully A sub-scale model with The Sub-Scale test was launch and recover a sub-scale Test comparable weights, lengths, and successful and has been 1.16 model of their rocket prior to masses will be launched prior to completed. See the CDR’s CDR. the CDR. “Sub-Scale” Flight section. All teams shall successfully Launch both complete and The project schedule will ensure a launch and recover their full- successful on February 18th. 1.17 Test full-scale rocket launch occurs scale rocket prior to FRR in its See “Full Scale Flight” for more before the FRR. final flight con- figuration. detail. The rocket has 3 bolts holding the nosecone to the bow body tube. These are located bow of Any structural protuberance on No structural protuberances will Test the burnout center of gravity but 1.18 the rocket shall be located aft of exist bow of the burnout center of Analysis has been cleared by NASA. No the burnout center of gravity. gravity. other structural protuberances exist bow of the burnout center of gravity.
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NASA Requirements Handbook Verification Status & Location Summarized Requirement Description of Verification Plan Number Method(s) The launch vehicle has The launch vehicle will follow all followed all prohibitions laid prohibitions laid out in section out in section 1.19 of the 2017 1.19 Vehicle Prohibitions Inspection 1.19 of the 2017 SL NASA SL NASA Student Handbook. Student Handbook. See “Vehicle Criteria” for full design. Vehicle must deploy a drogue Full-scale test flights resulted in Dual-deployment altimeters are parachute at apogee, followed by successful recovery events. See 2.1 Test programmed to fire ejection a main parachute at a much “Full Scale Flight” & charges at apogee and at 750 feet. lower altitude. “Recovery” for more. Sub-scale and full-scale test A successful ground ejection test Multiple ejection tests conducted ejections were successful – 8 for both parachutes must be 2.2 Test prior to sub- and full-scale consecutive full scale test conducted prior to sub- and full- launches. ejections. See “Ejection scale launches. Testing” section. Full-scale test flights resulted in No part of the launch vehicle Parachute sizes are optimized to Analysis kinetic energy below the 2.3 may have a kinetic energy minimize kinetic energy at Demonstration maximum allowable. See greater than 75 ft-lb at landing. ground impact. f “Recovery” section. Coupling tube constructed Recovery electrical circuits must Recovery electronics are housed completely independent of other 2.4 be independent of payload Inspection in a separate compartment. electronics. See “Recovery” circuits. section for more detail.
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NASA Requirements Handbook Verification Status & Location Summarized Requirement Description of Verification Plan Number Method(s) Redundant ejection charges Recovery system must include observed for each recovery Two PerfectFlite Stratologger CF 2.5 redundant, commercial Inspection event during full-scale test altimeters will be used. altimeters. flights. See “Recovery” section. Black powder ejection charges Motor ejection cannot be used Demonstration Black powder ejection charges successfully triggered recovery 2.6 for primary or secondary Inspection are used to deploy parachutes. events for full-scale test flights. deployment. See “Recovery” section. Rotary switches successfully Each altimeter must be armed by A separate switch accessible armed from rocket exterior for 2.7 a dedicated switch accessible Inspection through pressure sampling holes full-scale test flights. See from the rocket exterior. is used to arm each altimeter. “Recovery” section. Recovery altimeters were Separate 9-Volt batteries are Each altimeter must have a powered up for duration of each 2.8 Inspection wired to the power leads of each dedicated power supply. full-scale test flights. See altimeter. “Recovery” section. Recovery altimeters were Locking rotary switches are wired Each arming switch must be powered up for duration of each 2.9 Inspection to the switch leads of each lockable to the “ON” position. full-scale test flights. See altimeter. “Recovery” section. Pins sheared successfully Removable shear pins must be Three #2 nylon shear pins are during ejection testing and full- 2.10 used to seal the parachute Inspection used to seal each parachute scale test flights. See compartments. compartments. “Recovery” section.
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NASA Requirements Handbook Verification Status & Location Summarized Requirement Description of Verification Plan Number Method(s) All sections of launch vehicle remained tethered during full- Tracking device(s) must transmit Test All parts of the launch vehicle are scale test flights. Position data the position of any parts of the tethered together; position will be 2.11 Demonstration was successfully transmitted launch vehicle to a ground transmitted via a flight computer throughout each flight. See receiver. Inspection in the nosecone. “Line Item Budget” for exact GPS specifications. Recovery system electronics Recovery altimeter data showed must not be adversely affected Test Recovery electronics located in no signs of interference after 2.12 by any other on-board Inspection separate compartment. full-scale test flights. See electronics. “Recovery” section. Math model is used to develop Full scale flights resulted in safe Design container capable of spring system in conjunction with return of an egg – with ability to a concentric cylinder model to adjust to multiple eggs. See the 3.4.1 protecting an unknown object of Testing provide sufficient vibration “Payload Testing” section for unknown size and shape. dampening and force reduction. more detail, including a % reduction in force. The spring and concentric cylinder design will be tested Full scale flights resulted in safe with a matrix of different support return of an egg. See the Object must survive duration of 3.4.1.2 Testing materials as well as testing “Payload Testing” section for flight materials to assure the unknown more detail, including a % object(s) can survive the flight reduction in force. during demonstration.
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NASA Requirements Handbook Verification Status & Location Summarized Requirement Description of Verification Plan Number Method(s) Using only material already in Once the object is obtained, it the rocket, this setup was tested must be sealed in its housing During full scale flight, verify on February 18th for the full 3.4.1.4 until after the launch and no Test that an object can be contained scale flight and passed. See excess material may be added using no excess material. “Payload Testing” section for after receiving the object. more. Final assembly and pre-launch checklists will be created and Launch and safety checklist Each team shall use a launch and Inspection reviewed at the appropriate time used for full-scale test flight. 4.1 safety checklist Demonstration to ensure safe launch of the rocket See “Launch Operations” and all members involved in the section for more detail. launch Each team shall identify a The team has appointed a safety Safety officer Bryan Bauer student safety officer who shall officer to monitor the safety of oversaw both fabrication and 4.2 be responsible for the safety of Inspection the team throughout the project testing phases to ensure safe the team and ensure all proper and ensure all federal rules and and successful operations. rules and guidelines are followed laws are met. The team safety officer shall monitor team activities with an The team safety officer will Safety officer has monitored the emphasis on safety throughout monitor the progress of the full-scale testing, fabrication 4.3 the design, construction, and Inspection project emphasizing the proper and launch in order to ensure testing of the rocket by safety procedures for the current safe operations. maintaining MSDS sheets and stage of the project. hazard analyses.
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NASA Requirements Handbook Verification Status & Location Summarized Requirement Description of Verification Plan Number Method(s) The team has assigned an school faculty member to mentor the Each team shall appoint a mentor Dr. David Unger is the team project to provide valuable insight 4.4 who has certification and is in Inspection mentor; his information can be on the rocket design and good standing with the NRA. found on the cover page. construction as well as assume full liability of the rocket. Team is in compliance will all Team will converse with RSO at During test flights, teams shall rules and regulations set forth local rocketry club to ensure all of 4.5 abide by the rules and guidance Demonstration by local rocketry club their chapter’s rules and of the local rocketry club's RSO “BluesRocks”. See “Full Scale regulations are abided by. Flight” for more detail. Team will converse with NASA Team is in compliance will all lead safety officer and thoroughly rules and regulations set forth Teams shall abide by all rules set research all rules and regulations 4.6 Demonstration by FAA and NASA. See “Full forth by the FAA set forth by the FAA to ensure all Scale Flight” for more detail on rules and regulations are abided the flight. by. The team will continuously Students shall do 100% of the Demonstration demonstrate an independently The team has only used mentors 5.1 project excluding motor / black managed and executed project. for guidance and will continue Inspection powder handling. The team lead will routinely to do so. monitor this quality.
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NASA Requirements Handbook Verification Status & Location Summarized Requirement Description of Verification Plan Number Method(s) Documents for scheduling, Project plan is updated and can A detailed project plan shall be budget tracking, outreach, and 5.2 Demonstration be found in the “Project Plan” maintained. safety will be continuously section. updated and reported. The team lead will ensure that Foreign National members have Foreign National members shall 5.3 Inspection any Foreign National members been identified in emails with be identified by the PDR. are clearly indicated in the PDR. NASA. It will be checked that a list of All team members attending Team members have been team members, with indications 5.4 launch week shall be identified Inspection identified in emails with NASA, of those attending launch week, by the CDR. along with completed waivers. will be included in the CDR. The Educational Engagement Team has completed outreach The educational engagement lead shall confirm that all activities with over 200 students 5.5 requirement shall be met by the Inspection documentation has been received reached. Educational FRR. and approved by NASA prior to engagement is not discussed in the FRR. this report. Team members will periodically The team shall develop and host confirm that the website is Website has been developed 5.6 Test a website for documentation. functioning as intended by and is being updated. opening each posted document. All reports, presentation slides The team shall post & make The team lead shall confirm that and flysheets have been and available for download all 5.7 Inspection all documents are posted prior to will continue to be posted to the deliverables by the specified the specified date. team website by the deadline set date. forth by NASA.
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NASA Requirements Handbook Verification Status & Location Summarized Requirement Description of Verification Plan Number Method(s) The team lead shall confirm that All deliverables to the team All deliverables must be in PDF 5.8 Inspection all documents posted are in PDF website are upload in PDF format. format. format. The team lead shall ensure that a A table of contents must be See Table of Contents, Figures, 5.9 Inspection table of contents is located at the included in all reports. and Tables sections. start of each report. Page numbers shall be checked to Page numbers shall be provided See lower right hand corner of 5.10 Test the table of contents to ensure in each report. each report. continuity throughout the report. The team shall provide Demonstration Videoconference rooms will be Requirement met, same setup 5.11 videoconference equipment reserved and trialed immediately will be used for all future Test needed for reviews. prior to each design review. correspondence. The 12’ 1515 Rail used for sub- All teams shall use launch pads The team shall design the rocket scale launch operated as 5.12 Demonstration provided by the SLS provider. to utilize 1515 12’ launch rail. intended, see Full Scale Flight section. If software or applications are created (not planned) the team Software not designed by team. The team must implement the will abide by 36 CFR Part 1194. 5.13 Demonstration See “Line Item Budget” for EIT accessibility standards. Otherwise, all components exact electronic components. containing software will be checked to ensure compliance.
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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 Verification Description of Verification Status & Location Number Requirement Method Method Reports shall be completed, The team has completed all All reports shall be compiled at according to team schedule, prior reports on time. The dates can 1 least three days prior to NASA Demonstration to NASA due dates to allow for be seen in the “Schedule” due dates. revision time and mitigate risk of portion of the report. late submissions. At each team meeting, every sub- This has been maintained. It section lead will review the status was recently demonstrated at Each member of the team shall of their section with the entire the full-scale launch where 2 have a working knowledge of Inspection team. The team leader will team members had to work on each subsystem. confirm that the information each other’s sections. See “Full presented is sufficient. Scale Flight” for more details. Safety officer has asked 17 The safety officer will team members what the most Safety shall be made the team’s periodically ask team members important part of the project is 3 Demonstration first priority. what the most important aspect of and has had 15 “safety” the project is. answers. The two outliers have been reminded of safety. All altimeters shall be flown on Altimeters have all been Altimeters shall be in good sub-scale and full scale flight extensively tested and have 4 Test working order. tests. Altitude readings will be passed all tests. See “Altimeter compared to confirm consistency. Testing” section for more detail.
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Team Requirements Verification Description of Verification Status & Location Number Requirement Method Method The tracking system performed The tracking system shall be perfectly in the sub-scale & flown on the sub-scale and full The tracking system shall be in full-scale test. During the full 5 Test scale flight tests. This will be good working order. scale test, the tracking system used to find the rockets thus located the rocket over 1 mile confirming its operation. away. All altimeters will be triggered A solid output signal must be Test while voltage is read on the The output voltage is seen in 6 given from triggered altimeters. Analysis output. This output will be read real time at the base station. to confirm it is acceptable. This was completed for both All circuits shall be checked All circuits will be confirmed at sub-scale and full-scale tests. 7 Inspection prior to use. each node to ensure connections. Continuity and amperage were both inspected. The main parachute shall have an apparatus (strain gauge) attached Parachute force testing Impulse for the parachute Test to it that enables a force to be completed using acceleration 8 deployment shall be determined read as it opens at high speed. data on altimeter. See Analysis experimentally. This will cut down in the large “Parachute Deployment Force ambiguity that exists in Testing” section. estimating an impulse value.
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Team Requirements Verification Description of Verification Status & Location Number Requirement Method Method This testing has been bundled The spring constant shall be into the parachute deployment determined using forces related to force test. The parachute as it A spring constant for parachute what is experienced with Test relates to the body had its 9 cords shall be determined parachute opening. This helps Analysis acceleration measured, which as experimentally. when estimating energy a system includes the cords’ absorption by the cord when the expansion. See “Parachute chute opens. Deployment Force Testing”. From the mathematical model, This testing has been completed appropriate springs will be and selected springs were also Payload must reduce force felt 10 Testing selected to induce oscillation and tested to assure spring constants by object(s) by 50 % reduce force. These will be tested given by the manufacturer were by Charpy Impact Tests. accurate. The test to deduce the percent reduction in force and From the mathematical model, acceleration was completed appropriate springs will be however a faulty accelerometer Payload must reduce selected to insure acceleration 11 Testing caused data to be useless and acceleration of object(s) by 35 % graphs show 35 percent reduction therefore percent reduction from inputs. Will be tested via cannot be found. See Payload Charpy Impact Test. Testing section for full explanation.
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Team Requirements Verification Description of Verification Status & Location Number Requirement Method Method First, temperature sensitive 2 Full Scale ejection tests done Electronics must operate in cold Demonstration components will be identified. in cold (-2°C) environment, test 12 temperatures Testing Then components will be tested passed. See “Ejection Testing” in the cold with ejection testing. section. From simulations, the motor and Full-Scale completed and Mach Number will be less than Simulation aerodynamics of the rocket will simulations ran. See section 13 0.6 Test ensure the rocket has a Mach “Flight Simulations and number of 0.53 Altitude Predictions” From the simulations, the rocket Full-Scale completed and The rail exit velocity will be weight and motor section will simulations ran. See section 14 Simulation above 60 ft/s ensure of having a rail exit “Flight Simulations and velocity of 66.9 ft/s Altitude Predictions” From hand calculations to obtain Combustion analysis complete Complete a Combustion the temperature and pressures to and has a pressure of 2155 kPa. 15 Analysis on the Motor to obtain Simulation run the FEA analysis on the Located in Combustion Pressure of fuel ignition motor casing to find the Factor of Analysis in CDR. Safety of the motor casing Used hand calculations to Complete a Modal Analysis on Modal analysis complete with determine the natural frequency the Motor Mount System to an operating frequency not near 16 Simulation of the motor mount and then used ensure safety and stability of the natural frequency. See “Modal Finite Element Analysis to find rocket Analysis” in CDR. operational frequency
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Team Requirements Verification Description of Verification Status & Location Number Requirement Method Method Used hand calculations to have a Complete a Shear Stress verification of the results found Analysis complete with F.O.S. Analysis on the motor mount to 17 Demonstration using the Finite Element Analysis of 17.07. See “Shear Stress ensure that the epoxy being used to find the Factor of Safety of the Analysis” in CDR. will withstand the motor forces motor mount Calculated the Factor of Safety of Have a Factor of Safety above 2 Combustion and Shear Stress Complete with F.O.S. of 103.6 18 for the Combustion Analysis and Demonstration areas. Found Combustion factor & 17.07, respectively. See Shear Stress Analysis of Safety to be 103. Found Shear “Propulsion” Section in CDR. Stress factor of safety to be 17 Use OpenRocket and Rocksim to Full scale test apogee of 5,291 Reach an altitude between 5,200 Simulation simulate the altitude of the full- 19 feet. See “Full Scale Flight” and 5,400 feet Testing scale rocket. Test the full-scale to section for more detail. see the actual altitude Using simulation and Ballast adjusted on February demonstration of design, the team Design flexibility on full-scale 18th to change height for three Demonstration will prove that on test launch day, 20 test launch day to raise or lower different flights. See “Full Simulation small changes can be made to altitude on a second test-flight. Scale Flight” for more detail on raise or lower altitude for a the configurations. second flight.
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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
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
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.
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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
Due Date Activity 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 Plan Duration Actual Start % (Planned) Actual (beyond plan) % (Unplanned)
T/M RESPONSIBLE PLAN ACTUAL ACTUAL (Week 1 ends September 4th, 2016) ACTIVITY PLAN START PERCENT COMPLETE DURATION START DURATION PERIODS 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 35 100% Proposal David 1 4 1 4 100% Preliminary Design Report David 6 4 6 4 100% PDR Presentation David 8 2 8 2 100% Interim Design Report David 12 3 13 2 100% Critical Design Report David 11 5 10 8 100% CDR Presentation David 15 5 15 5 100%
Flight Readiness Report David 23 4 24 3 Reporting 100% FRR Presentation David 26 2 26 2 0% Project Final Report David 31 2 0% Launch Readiness Review David 29 4 0% Post Launch Assesment David 33 2 100% Budget Creation David 1 1 1 5 100% Website Creation Bryan 1 3 2 2 100% Motor Type Selection Andrew G 1 3 1 3 100% Motor Mount Design Andrew G 1 5 1 5 100% Rocksim Model Andrew G 3 18 3 4 100% Body Component Selection Torsten 1 6 1 6 100% 3D Rocket Model Torsten 4 11 4 6 100% CFD Model Torsten 15 6 13 2 100%
Payload A Design Justin 1 9 1 9 Design 100% Payload B Design Braden 1 11 1 11 100% Data Acquisition Design David 3 6 3 2 100% Data Transmission Design David 3 6 3 2 100% Design of Recovery System Andrew S 1 9 1 9 100% Design Tracking System Andrew S 9 4 9 4 100% Design Education Activity Bryan 1 8 1 12 100% Propulsion Construction Andrew G 10 6 12 4 100% Body Construction Torsten 11 11 10 13 100% Payload A Construction Justin 9 14 9 14 100% Payload B Construction Braden 12 11 12 12
Construction 100% Recovery System Construction Andrew S 9 12 9 13 100% Data Systems Construction David 8 13 9 13 100% Scale Model Construction Torsten 12 3 11 4 100% Scale Model Test Team 14 2 14 2 100% Bulkhead Testing Rakan 22 3 24 2 100% Payload Testing Braden 14 9 20 6
Testing 100% Parachute Testing Andrew S 23 6 25 1 30% Wind Tunnel Testing Feras 23 9 25 3 100% Recovery Testing Andrew S 19 7 22 4 100% Educational Engagement Bryan Ongoing Ongoing 0% Preparation for Competition David 31 1 0% Competition David 32 1
Figure 52 - Gantt Chart
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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 Aerodynamics 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 Propulsion 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
Recovery 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
ElectronicPayload 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
Main Payload Main 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
Safety / Educational Engagement $ 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 Administrative / Travel / Administrative 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 Launch Pad Launch $ 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
ScaleModel 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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Appendix G – Task Breakdown Task Breakdown
Task* Responsible Start Date End Date Comments Person Estimated Actual Estimated Actual
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
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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
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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
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Appendix H – Electrical Diagrams Electrical Diagrams
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Appendix I – Payload Accelerometer Graphs Payload Accelerometer Graphs
90 Degree Cotton Fill Large Bulb 30000
20000
10000
0 Ax 0 20 40 60 80 100 120 Ay -10000
Acceleration Az -20000
-30000 -
-40000 Time Step (s) 31727
Figure 72 – 90 Degree Cotton Fill Large Bulb
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90 Degree Paper Fill Large Bulb 3000
2000
1000
0 Ax 0 10 20 30 40 50 60 Ay -1000 Az
-2000
-3000 -3434
-4000
Figure 73 – 90 Degree Paper Fill Large Bulb
90 Degrees Packing Peanuts Large Bulb 40000 32185 30000
20000
10000 Ax 0 Ay 0 20 40 60 80 100 120 140 Az -10000
-20000
-30000
-40000
Figure 74 – 90 Degrees Packing Peanuts Large Bulb
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90 Degrees Large Bulb Only (no fill) 3000 2526
2000
1000 Ax 0 Ay 0 1 2 3 4 5 6 Az -1000
-2000
-3000
Figure 75 – 90 Degrees Large Bulb Only
90 Degrees DogBrag Fill Large Bulb 4000 3249 3000
2000
1000 Ax 0 Ay 0 5 10 15 20 25 30 35 Az -1000
-2000
-3000
-4000
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Figure 76 – 90 Degrees DogBrag Fill Large Bulb
90 Degrees Base Value 15000 -18895 10000
5000
0 Ax 0 1 2 3 4 5 6 7 8 9 -5000 Ay Az -10000
-15000
-20000
-25000
Figure 77 – 90 Degrees Base Value
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Appendix J – Wind Tunnel Uncertainty Wind Tunnel Uncertainty
Sample calculation of drag coefficient’s uncertainty:
Assuming density (휌) is constant.
휖∗푤∗푡2 퐶퐷 = 푈2 6∗퐿∗휌∗ ∗퐴 2 푐
휕퐶퐷 퐸∗푤∗푡2 = 푈2 휕휖 6∗퐿∗휌∗ ∗퐴 2 푐
휕퐶퐷 휖∗푤∗푡2 = 푈2 휕퐸 6∗퐿∗휌∗ ∗퐴 2 푐
휕퐶퐷 2∗퐸∗푤∗푡 = 푈2 휕푡 6∗퐿∗휌∗ ∗퐴 2 푐
휕퐶퐷 휖∗퐸∗푤∗푡2 = −( 푈2 ) 휕퐿 6∗퐿2∗휌∗ ∗퐴 2 푐
휕퐶퐷 2 ∗ 휖 ∗ 퐸 ∗ 푤 ∗ 푡2 = −( ) 휕푈 푈3 6 ∗ 퐿 ∗ 휌 ∗ 2 ∗ 퐴푐
휕퐶퐷 휖 ∗ 퐸 ∗ 푤 ∗ 푡2 = −( 2 ) 휕푈 푈 2 6 ∗ 퐿 ∗ 휌 ∗ 2 ∗ 퐴푐
For velocity:
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휕푈 2 = √ 휕∆푃 2 0.00226 ∆푃 2 (푙푏푓⁄푓푡 ) (푙푏푓⁄푓푡 )
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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.
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