University of Evansville
Student Launch
Enclosed: Preliminary Design Review
Submitted by: 2016 – 2017 Rocket Team Project Lead: David Eilken
Submission Date: November 04, 2016
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 ...... iv
List of Tables ...... vi
PDR 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
Selection, Design, & Rationale of Launch Vehicle ...... 4
Mission statement ...... 4
Mission Success Criteria ...... 4
System Level Alternatives and Analysis ...... 6
Component Alternatives ...... 12
Motor Alternatives ...... 22
Recovery...... 26
Payload ...... 32
Electronic Payload ...... 32
Fragile Material Payload ...... 34
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Mission Performance Predictions ...... 45
Safety ...... 53
Overview ...... 53
Final Assembly Checklist ...... 55
Launch Procedures Checklist ...... 57
Personnel Hazard Analysis...... 59
Failure Modes and Effects Analysis...... 60
Environmental Considerations ...... 61
General Risk Assessment ...... 63
Project Plan ...... 64
Requirements Compliance ...... 64
Budget ...... 75
Schedule ...... 76
References ...... 79
Appendix A – Machine Prints...... 80
Appendix B – OpenRocket Simulation...... 87
Appendix C – Parts List ...... 91
Appendix D – Task Breakdown ...... 93
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List of Figures
Figure 1 - Updated 3D Model of Launch Vehicle ...... 2
Figure 2 - Rocket System Decomposition ...... 6
Figure 3 - Weight breakdown (all weights are in lbf) ...... 7
Figure 4 - Dimensioned drawing of full body (all dimensions in inches) ...... 8
Figure 5 - Subsection dimensions ...... 8
Figure 6 - Nosecone mounting diagram...... 9
Figure 7 - Exploded View of the Motor Mount ...... 24
Figure 8 - Propulsion Components Labeled ...... 25
Figure 9 - Dimensional Drawing for the Motor Mount ...... 25
Figure 10 - PerfectFlite Stratologger CF Altimeter ...... 27
Figure 11 - Block diagram of major recovery system electrical components ...... 28
Figure 12 - Recovery bay bulkheads and hardware ...... 29
Figure 13 – Exploded View; Recovery System ...... 30
Figure 14 - Recovery system layout within airframe ...... 30
Figure 15 – Tethering of Rocket Sections ...... 31
Figure 16 - Electronic Payload within Nosecone ...... 32
Figure 17 - Exploded View of Electronic Payload ...... 32
Figure 18 - Exploded Electronic Payload View with Nosecone ...... 33
Figure 19 - Top View, Assembled Electronic Payload ...... 33
Figure 20 - Bottom View, Assembled Electronic Payload ...... 33
Figure 21 - Payload Exploded View ...... 35
Figure 22 - Components of the Main Payload ...... 36
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Figure 23 – Payload Inner Cylinder ...... 37
Figure 24 - Springs Attached to the Main Payload and Bulkhead Attachment ...... 38
Figure 25 - System Drawing and Force Balance ...... 40
Figure 26 - Free Body Diagram for the Entire Rocket and Force Balance ...... 41
Figure 27 - Free Body Diagram for the Spring Assembly and Force Balance ...... 41
Figure 28 - Free Body Diagram for Cylinder 1 and the Fragile Material and Force Balance 42
Figure 29 – Simulink Mathematical Model ...... 44
Figure 30 - Predicted Altitude from OpenRocket Simulation ...... 47
Figure 31 - OpenRocket Flight Simulation Inputs ...... 48
Figure 32 - Predicted Altitude from Rocksim Simulation ...... 49
Figure 33 - Inputs for Rocksim Simulation ...... 50
Figure 34 - Thrust Curve from AeroTech Motor ...... 50
Figure 35 - Thrust Curve for the L850W Motor in OpenRocket ...... 51
Figure 36 - Thrust Curve for the L850W Motor in Rocksim ...... 51
Figure 37 - Center of pressure and gravity ...... 52
Figure 38 - Gantt Chart ...... 77
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List of Tables
Table 1 - Decision Matrix Key ...... 12
Table 2 - Decision Matrix: Body Tube ...... 12
Table 3 - Decision Matrix: Fin and Nosecone Material ...... 13
Table 4 - Decision Matrix: Bulkhead Material ...... 14
Table 5 - Decision Matrix: Fin Shape ...... 14
Table 6 - Decision Matrix: Nosecone Shape ...... 15
Table 7 - Decision Matrix: Motor Mount Design ...... 16
Table 8 - Decision Matrix: Centering Rings ...... 17
Table 9 - Decision Matrix: Recovery Altimeter ...... 19
Table 10 - Decision Matrix: Recovery Harness Material ...... 20
Table 11 - Decision Matrix: Drogue Parachute ...... 21
Table 12 - Decision Matrix: Main Parachute ...... 22
Table 13 – Motor Considerations and Specifications ...... 23
Table 14 - Testing Matrix for Fragile Material ...... 39
Table 15 - Force Events for the Simulink Model ...... 42
Table 16 - Final Values for Constants ...... 45
Table 17 - Kinetic energy of each section upon landing ...... 52
Table 18 - Landing site distance from launch site by wind speed ...... 53
Table 19 - Personnel Hazard Analysis ...... 59
Table 20 - Failure Modes and Effects Analysis ...... 60
Table 21 - Environmental Consideration Analysis ...... 61
Table 22 - General Risks Associated with the Project ...... 63
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Table 23 - Requirement Compliance ...... 64
Table 24 - Team Requirements ...... 73
Table 25 - Section Level Budget ...... 76
Table 26 - Funding Sources ...... 76
Table 27 - Critical Dates ...... 78
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PDR Summary
Project ACE plans to field a 111” long, 35-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 waterproof 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. This 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 will be contained, will feature 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 down the rocket from the payload is the recovery system. This system features completely redundant separation circuits. At apogee, a 48” drogue chute will eject, followed by a 96” main chute closer to ground level. At the aft end of the rocket sits the propulsion section.
A 75-mm L-850W Aerotech motor will propel the rocket for just over four seconds. This motor will be held in place via 6061-T6 Aluminum centering rings and thrust plates. All components will be housed in two carbon fiber body tubes. The fins, which adhere to the centering rings and body tubes, will be 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 bow body tube was elongated by 8” to accommodate the design changes made to the main payload. The fin thickness was also decreased to 0.125” and designed to have a beveled leading edge. This will decrease the drag on the launch vehicle. Lastly, it was determined through manufacturer specifications that the exact length of the nosecone will be 21.75”. The remainder of the vehicle criteria remained unchanged. An updated 3D model of the launch vehicle can be seen in Figure 1.
Main Payload
Figure 1 - Updated 3D Model of Launch Vehicle Changes Made to Payload Criteria
The spring system used to support the payload added 5 base springs after the math model proved that wire rope isolators alone would not be sufficient. To accommodate this design change, the entire previous payload was re-designed to oscillate within the body tube. The spring selections originally planned also changed due to system optimization through a math model.
More detail can be found in the payload section.
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Changes Made to Project Plan
Few changes have occurred to the project plan since the proposal was submitted. NASA SLI officials have indicated that the due date for PDR documentation has been moved to November
4th, 2016 (originally October 28th). Despite this, Project ACE has decided to keep to the schedule of having PDR documents completed by October 26th. This will enable the team to focus on the build phase of the sub-scale rocket. More on the schedule can be found in the “Schedule” section of this report.
The budget has been decreased by $350.00. Additionally, funding has been allocated in a slightly different fashion than in the proposal. The reason for this is twofold: first, the motor had unforeseen hardware costs associated with it, increasing the funds needed for that section. The travel and lodging portion of the budget decreased substantially, as Project ACE decided not to have the team cover any meal costs. Also, it was determined that advisor expenses would come out of the University of Evansville College of Engineering and Computer Science budget instead of the project budget. A detailed budget breakdown can be found in the “Budget” section of this report.
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Vehicle Criteria
Selection, Design, & Rationale of Launch Vehicle
Mission statement
Project ACE is an interdisciplinary university project with the united goal of constructing and flying a high powered aircraft with a unique experimental payload. Our team intends to perform at a high level at the national competition and pass down the knowledge gained from this experience to current underclassmen and future Project ACE members.
Mission Success Criteria
1. Aerodynamics
a. The airframe, nose cone, and fins should remain intact for the duration of the
flight.
b. The airframe, nose cone, and fins should be reusable for any following flights.
c. The airframe and nose cone should protect all internal components from
damage from external sources.
2. Propulsion
a. The vehicle should attain an apogee between 5,125 feet and 5,375 feet.
b. The vehicle should remain below Mach 1.
c. The motor mount should withstand propulsion forces and remain reusable for
any following flights.
3. Recovery
a. The drogue parachute and main parachute are ejected at apogee and 1000 feet,
respectively.
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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 should be able to be received
remotely during and after the vehicle’s flight.
b. The electronic payload should withstand flight forces and remain reusable for
any following flights.
c. The electronic payload should accurately determine the apogee of the rocket.
5. Main Payload
a. The fragile object(s) should remain undamaged.
b. The force felt by the payload should be reduced by 50% for each of the areas
of interest: takeoff (thrust curve, parachute deployment, and landing.)
c. The force felt by the payload should be reduced by 35% for each of the areas
of interest: (thrust curve, parachute deployment, and landing.)
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System Level Alternatives and Analysis
The launch vehicle was designed with 5 interconnected systems: the airframe, electronic payload, main payload, recovery, and propulsion. These systems and relationships can be seen in
Figure 2. The airframe is the parent system and houses all the sub-sections.
Figure 2 - Rocket System Decomposition
The full weight of the launch vehicle is 35.19 lbf. A weight breakdown of the rocket and the individual subsections can be seen Figure 3.
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Figure 3 - Weight breakdown (all weights are in lbf) The purpose of the airframe is to provide a structure for the internal systems and protect them from external stresses. The airframe was designed to be comprised of two carbon fiber body tubes and an ogive fiberglass nosecone. The body tubes will be made of carbon fiber. Both body tubes will have a diameter of 5.5”. The aft body tube will have a length of 48”. The bow body tube will have a length of 41”. The nosecone will be made of fiberglass and will have a 4:1 ogive profile. The total length of the nosecone is 21.75”. A dimensioned drawing of the full body is provided in Figure 4.
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Figure 4 - Dimensioned drawing of full body (all dimensions in inches) The airframe will house 4 main systems: electronic payload, main payload, recovery, and propulsion. The allocated space and sizing for the individual subsections can be seen below in
Figure 5.
Figure 5 - Subsection dimensions The two body tube system was chosen over a single body tube system. This was done in order to incorporate a dual deployment recovery system that would separate between the two body tubes. The retention system for the nosecone is currently designed to be mounted with 3 bolts and 3 adhesive mount nuts. This was chosen over alternatives such as a threaded rod mounted down the length of the nosecone or threads on the interior wall of the bow body tube.
The current system was chosen because it allows the bow body tube to remain completely free of permanent mounting hardware. This allows the main payload to be removed and inserted with ease. This design can be seen below in Figure 6.
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Figure 6 - Nosecone mounting diagram The purpose of the electronic payload is to provide an official altitude, GPS coordinates for the launch vehicle, and hold ballasts. It will be mounted with a gasket and removable mount combination in the nosecone with the electronics facing towards the bow end of the rocket. This will provide an added measure of security towards water damage.
The Atlus TeleMega was chosen against other altimeters because it records flight data in addition to apogee and GPS location. Much of this data can be compared to RockSim. Other altimeters were cheaper, however, the extra data (such as rocket tilt) was determined to be worth the cost difference.
The purpose of the main payload is to protect fragile materials. It consists of a concentric cylinder design as well as a series and parallel spring system. The inner cylinder utilizes wire rope isolators to absorb smaller vibrations while larger springs at the base of the cylinder reduce the force of large impulse impacts such as takeoff, landing, and main parachute deployment. 9 | P a g e
Prior to choosing the main payload design that currently exists, several options were discussed. One option was simply a payload bay with support material and a cap that had built in damping to hold the unknown fragile material object(s) in place and hopefully protect them.
The other alternative was the concentric cylinder design with wire rope isolators; however, the math model used to predict the behavior of the system showed this was not sufficient. That is what prompted the additional larger springs that were added in series with the wire rope isolators.
The recovery system serves to return the launch vehicle to the ground safely, minimizing the ground impact velocity to preserve the structural components of the rocket as well as the fragile payload. A dual-deployment system utilizing a 36" drogue parachute and a 96" main parachute has been designed to use identical, redundant electrical systems to trigger black powder ejection charges. The electrical systems will be housed in a coupling tube that unites the bow and aft body tubes. The drogue chute will be packed in the bow tube, and the main chute in the aft tube.
All sections of the rocket will be tethered together using a tubular nylon recovery harness.
Several system-level alternatives were considered for the recovery system. In particular, a gas ejection system was investigated, in which a canister of compressed CO2 is used to pressurize the parachute compartment during a deployment event. While gas ejection systems do not subject the parachute to the high temperatures of a black powder ejection, they tend to be heavier, more complicated, and more expensive than a simple black powder ejection. For these reasons, a gas ejection system was not selected.
Additionally, different parachute deployment schemes were considered. In many rockets, the drogue parachute is packed underneath the nose cone and deployed by blowing the nose cone out
10 | P a g e of its body tube. This method was not selected because it would require that the recovery electronics be located in close proximity to the transmitting components of the competition altimeter, which could create unwanted interference. Recovering the rocket in multiple components was also considered; for example, the bow and aft body tubes could be tethered together after drogue deployment and split during the main deployment to be descended under separate main parachutes. This setup was not selected due to limitations on body tube space created by the main payload.
Lastly, the aft body tube houses the propulsion section. The purpose of the propulsion section is to propel the launch vehicle to a height of 5,280 ft. The propulsion section was designed to house 3 centering rings and an engine block (all made of 0.25” aluminum). The aft body tube will be slotted to allow the fiberglass fins to be attached to the inner tube and centering rings.
This adds further support for the fins and centering rings. The inner tube will be made of blue tube and have a 3.1” OD and 20” length. The inner tube will house an Aerotech L850W motor with a max thrust and impulse of 1185 newtons and 3695 newton seconds, respectively. The fins will have a clipped delta design.
The propulsion system was designed around a few key criteria. First, it was decided to use 3 centering rings versus 2 centering rings. This decision was made to increase stability of the inner tube. With a 3-centering ring system, two centering rings can support the fin tabs and one centering ring can be used as a thrust plate and serve as a mounting point for the motor retention system. Secondly, two motor retention systems were evaluated. The first system included threaded rods mounted to the engine block. The second system mounts directly to the furthest aft centering ring. The second system was chosen because of the decreased complexity and decreased weight.
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Component Alternatives
Decision matrices were used to visually and concisely evaluate multiple component-level options. These matrices can be seen throughout the report, and the key that they follow is located in Table 1. Bolded and underlined options indicate design selections. Discussion of the various decision matrices can be found immediately following each matrix.
Table 1 - Decision Matrix Key
Decision Matrix Criteria
О – Good Δ – OK X – No Good
Table 2 - Decision Matrix: Body Tube Decision Matrix – Body Tube Option Cost Strength Ductility Overview Decision Explanation Carbon Fiber provides the highest Carbon Fiber tensile strength and lowest X О О ductility at the highest cost Fiberglass provides a moderate Fiberglass strength, ductility, and cost Δ Δ Δ relative to Carbon Fiber and Blue Tube. Blue Tube provides the lowest Blue Tube ductility and strength at the О X X lowest cost.
Material considerations for the airframe included fiberglass, carbon fiber, and Blue Tube.
The team intends to use carbon fiber for the body tubes because it has a higher tensile strength, lower density, and a lower ductility compared to that of fiberglass or Blue Tube. Flexibility in a rocket airframe is an unwanted characteristic so a lower ductility is beneficial. In addition, the
12 | P a g e higher tensile strength of carbon fiber will ensure a higher allowable stress and a higher factor of safety than that of fiber glass.
Table 3 - Decision Matrix: Fin and Nosecone Material Decision Matrix – Fin and Nosecone Material Option Cost Strength Ductility Overview Decision Explanation Carbon Fiber provides a high Carbon Fiber tensile strength and low ductility X О О at a high cost. Fiberglass provides a moderate Fiberglass strength, ductility, and costs О Δ Δ significantly less than Carbon Fiber or ULTEM. ULTEM provides a high tensile ULTEM strength and low ductility at a X О О high cost.
The material for the fins and nosecone will be G-10 fiberglass because it is commercially available at a low cost. Carbon fiber and ULTEM plastic are also materials used for fin design; however, these did not provide adequate benefit to mitigate the significantly higher cost. This is because the nosecone and fins are not being required to undergo the same stresses caused by recovery process as the body tubes, so the additional strength of carbon fiber is not sufficient for these components.
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Table 4 - Decision Matrix: Bulkhead Material Decision Matrix – Bulkhead Material Option Cost Strength Weight Overview Decision Explanation Aluminum offers the highest Aluminum strength of all materials X О Δ considered. It comes at an increased cost and weight. Plywood offers the lowest cost Plywood and weight at the price of O X O strength. Fiberglass offers a moderate Fiberglass alternative to plywood and Δ Δ Δ fiberglass.
The bulkheads will be made of aluminum. Aluminum will be used to ensure the recovery and propulsion sections have strong attachment points. Fiberglass and plywood are common choices for bulkheads because they are sturdy, lightweight materials. However, since the design of the rocket is for an L-class motor, weight is not a significant constraint for material selection. This allows the team to choose the material with the highest tensile strength (aluminum) over fiberglass or plywood.
Table 5 - Decision Matrix: Fin Shape Decision Matrix – Fin Shape Ease of Likelihood Option Stability Overview Decision Explanation Manufacturing of Damage The Clipped Delta is the easiest to Clipped Delta manufacture and offers moderate Δ О O stability and drag. Trapezoidal The Trapezoidal offers the lowest X Δ О drag but the least stability. Tapered Swept The Tapered Swept offers the О Δ X highest drag but the least stability.
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The clipped delta design will be used for the fins. This design was chosen over other possible design choices such as a trapezoidal or tapered swept design. The difference between these designs is the sweep angle. This angle affects the center of pressure (CP) and thus affects stability. The clipped delta design was chosen after OpenRocket simulations and research was done on the various design choices. The research and simulations found the benefit of a different sweep angle to be minimal (<0.1 calipers stability increase). Additionally, changing the sweep angle to increase the stability would move the trailing edge of the fins aft of the end of the rocket. This would require the weight of the rocket to sit on the fins and increase the likelihood of damage.
Table 6 - Decision Matrix: Nosecone Shape Decision Matrix – Nosecone Shape Option Cost Drag Overview Decision Explanation The Ogive nosecone is the most difficult to Ogive manufacture and thus the most expensive but Δ О offers the lowest drag. Elliptical The Elliptical nosecone can be purchased at a O Δ moderate cost for a moderate drag. The Conical nosecone is the easiest to Conical manufacture and thus the least expensive but O X offers the highest drag.
Although the Ogive nosecone shape is the most difficult to manufacture, it offers the lowest drag of all nosecone profiles. For this reason, the nosecone will be purchased.
With the components of the body for the initial design of rocket chosen, the motor was the next area of the design. The first design of the motor was to use a cluster motor featuring three lower level motors to power the rocket. The other design consideration was using a single large motor to power the rocket.
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Table 7 - Decision Matrix: Motor Mount Design Decision Matrix – Motor Mount Design Against Option Cost Safety Overview Decision Explanation Regulations Cluster Motor The cluster motor would cost less and reach the optimal altitude Mount X O Δ with minimal safety concerns. Single Motor The single motor cost is high and creates concerns about safety and Mount X Δ O reaching altitude.
The motor mount that Project ACE was originally going to use was for a cluster motor configuration. This was due to the low cost of low level motors compared to a single large motor. Also, the cluster motors provide the ability to “mix and match” motors. The safety and complexity of the cluster motor, however, were concerns. There exists a heightened chance of misfires with use of more than one motor. There is also a chance that one motor does not ignite with the initial light, but could light from the other motors which is a clear safety concern. Table
7 shows the decision matrix for the motor mount design.
Originally, the single motor mount was the back-up plan. As previously mentioned cost was a major concern with the single, large motor design. From a first inspection, the cost for a single large motor was five times that of a cluster motor configuration. Additionally, few large motors were suitable to reach the one-mile mark. This, in turn, limited the design of the motor mount due to the lack of motor choices. The forces being produced with a single large motor may also be more concentrated within the mounting configuration, requiring more robust mounting. Table 7 shows the decisions for the motor mount.
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Initial designs for the motor mount were considered, before the 2017 handbook was posted for the USLI teams and utilized a cluster motor. Once the new handbook was released, the team learned that cluster motors had been disallowed. Thus, the team decided to go with the single motor and single motor mount for the propulsion for the rocket. The single motor mount design would use a larger motor and thus concerns arose about the shear forces being produced on the centering rings and the bulkhead. These concerns will be mitigated using FEA.
Table 8 - Decision Matrix: Centering Rings Decision Matrix – Centering Rings Option Cost Strength Weight Overview Decision Explanation Plywood is great for weight and Plywood X cost but the strength is a problem O О for large motors The cost of fiberglass is budget- G10 Fiberglass able because of the high strength Δ О О and the weight of the material Aluminum has a good cost Aluminum associated with machining it in Δ О X house with high strength. Only concern is the weight
The structural integrity of centering rings was already under review when the initial motor mount design was decided. This was due to the shear forces that could be expected with high power rocket motors. Due to this, strength was the major criterion that was used to select centering rings material. Table 8 shows the decision matrix for the centering rings. Plywood was the first material considered because of its low cost and low weight. However, the strength of the material (primarily Tensile Strength) was deemed significantly more important than cost or weight.
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The next material that was considered was G10 fiberglass. Fiberglass has high strength and low weight. This was appealing as the forces could likely be handled by the material and the low weight aided in raising the rockets altitude. However, the cost of Fiberglass is significantly greater than that of plywood.
The last material that was considered was 6061-T6 aluminum. This was researched due to the high strength and machinability of the material. The cost of the material is manageable, especially since all machining would be conducted by the team. The only problem with the aluminum is the weight.
Weight was decided to be a minor factor. Thus, the material that was selected was the aluminum. As it turned out, the added weight of the aluminum helped with controlling the altitude and bringing the rocket down to a desirable apogee. Also with the strength of aluminum being so great, the risk of the material shearing is low.
Several dual-deployment altimeters were considered for the recovery electronics system; the PerfectFlite Stratologger CF, the AltusMetrum EasyMini, and the Entacore AIM3. To select this component, cost was given priority, as two of the selected altimeter type would need to be purchased to create redundancy within the system. All altimeters considered had similar feature sets which were sufficient for the purposes of the rocket, as more complex data collection and transmitting functions will be handled by the competition altimeter in the nosecone. The
PerfectFlite Stratologger CF was selected. The decision matrix for the altimeter can be seen in
Table 9.
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Table 9 - Decision Matrix: Recovery Altimeter Decision Matrix – Recovery Altimeter Option Cost Feature Set Power Draw Overview Decision Explanation PerfectFlite For a low cost, this altimeter Stratologger provides a full set of features with O O Δ a higher power draw. CF
AltusMetrum For a medium cost, this altimeter provides a reduced feature set EasyMini Δ Δ О with a low power draw. Entacore For a high cost, this altimeter provides a full set of features with AIM3 X O О a low power draw.
Three materials are common when choosing a recovery harness for high-powered rockets: elastic, kevlar, and nylon. As this is a critical component, cost was not considered to be a high priority in the decision-making process. In order to reduce the maximum forces experienced by the rocket, a material with moderate elasticity was sought – high elasticity in the recovery harness can cause the tethered components to snap back and collide with one another.
The large forces involved with parachute deployment require a material with a high breaking strength. An elastic recovery harness would not be an acceptable selection due to its low strength and high elasticity. While Kevlar is incredibly strong, it has almost no elastic potential, which would do little to reduce the forces experienced by the rocket. Nylon was selected because it maintains a moderate degree of elasticity with a breaking strength well above the maximum force experienced by the rocket. Table 10 shows a decision matrix for the recovery harness material selection.
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Table 10 - Decision Matrix: Recovery Harness Material Decision Matrix – Recovery Harness Material Option Cost Strength Elasticity Overview Decision Explanation For a very low cost, elastic Elastic provides a low-strength, high- O X X elasticity solution. For a high cost, Kevlar provides Tubular Kevlar the greatest strength with a very X O Δ low elasticity. Tubular For a medium cost, nylon provides acceptable strength at a Nylon Δ Δ О medium elasticity.
After investigating many parachutes from multiple manufacturers, the field was narrowed to focus on three different diameter “Fruity Chutes” parachutes for each the drogue and the main.
Fruity Chutes was selected as a manufacturer based on a reputation for tough, well-made parachutes, as well as the small packing volume of their parachutes relative to their competitors’ products. In the selection of both parachutes, cost was deemed to be of minor importance due to the critical nature of the recovery system.
Drogue parachute selection focused primarily on ensuring that the initial descent rate is low enough to minimize the force of the main parachute inflation, while keeping the initial descent rate high enough to ensure that the main parachute inflates predictably. Simulations were conducted in OpenRocket for each parachute diameter. A 24” parachute caused the rocket to experience high accelerations during main parachute deployment, which could damage the rocket or the fragile material payload. A 48” parachute resulted in an initial descent rate that may not allow the main parachute to inflate properly. A 36” drogue parachute was selected to ensure
20 | P a g e that the main parachute inflates while limiting maximum acceleration. Table 11 shows a decision matrix for the drogue parachute selection.
Table 11 - Decision Matrix: Drogue Parachute Decision Matrix – Drogue Parachute Descent Option Cost Max Force Overview Decision Explanation Rate For a low cost, a relatively quick 24” Fruity descent rate can be achieved, but Chutes Classic O Δ X at the cost of a large maximum Elliptical force at main parachute deployment. 36” Fruity For a low cost, a good descent Chutes rate can be achieved with an acceptable maximum force at Classic O O Δ main parachute deployment. Elliptical 48” Fruity For a medium cost, a relatively Chutes Classic slow descent rate can be achieved Δ Δ О with a low maximum force at Elliptical main parachute deployment.
Main parachute selection focused on minimizing the kinetic energy of the rocket at ground impact, as this event has the greatest potential for causing costly damage to the rocket. Managing main parachute deployment acceleration was also a consideration. Using OpenRocket simulations, a 72” parachute resulted in ground impact kinetic energy greater than the 75 ft-lbf allowed by NASA. A 96” parachute was selected to give a maximum kinetic energy of 29.4 ft- lbf for the aft body tube, which is the heaviest section of the rocket.
Table 12 shows a decision matrix for the main parachute selection.
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Table 12 - Decision Matrix: Main Parachute Decision Matrix – Main Parachute Ground Option Cost Impact Max Force Overview Decision Explanation Velocity 72” Fruity For a low cost, an unacceptable Chutes Iris ground impact velocity can be O X O achieved with a good maximum Ultra force. 84” Fruity For a low cost, an acceptable Chutes Iris ground impact velocity can be O Δ Δ achieved with an acceptable Ultra maximum force. 96” Fruity For a medium cost, a good impact Chutes Iris velocity can be achieved with an Δ O Δ acceptable maximum force. Ultra
Motor Alternatives
The motor was decided to be either a K or L class motor upon running simulations in
OpenRocket. With the range of motors narrowed, 54mm motors were selected as that diameter was conducive to the rocket dimensions. The motors were then narrowed further by length.
Finally, simulations were run on each motor to see the apogee obtained and the final motors were selected by running multiple simulations. The final three motors that were considered were from
AeroTech, Aminal Motor Works, and Cesaroni Technology. The motor data can be found in
Table 13.
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Table 13 – Motor Considerations and Specifications
Manufacturer AeroTech Cesaroni Technology Inc Animal Motor Works
Make L850W L800 L1080BB-P Total Impulse 3695 Ns 3731 Ns 3686 Ns Weight 8.1 lbs 7.75 lbs 7.92 lbs Weight Empty 3.54 lbs 3.79 lbs 4.13 lbs Length 20.9 in 19.1 in 19.6 in Diameter 2.95 in 2.95 in 2.95 in Type Reloadable Reloadable Reloadable Burn Time 4.24 s 4.63 s 3.31 s Average Thrust 868 N 805 N 1112 N Max Thrust 1185 N 1024 N 1258 N Altitude Reached 5,379 ft 5,460 ft 5, 329 ft
The L850W motor from AeroTech was ultimately selected. Cesaroni was not producing motors at the time of selection and a strict time schedule needed to be kept for the project. The
L1080BB-P motor from Animal Motor Works was not chosen because of its relatively high empty weight. Using the L850W motor, the rocket has achieved a thrust to weight ratio of
5.61:1. The velocity that the rocket experiences (max) is 592 ft/s and an acceleration of 208 ft/s2.
The mach number for the rocket is 0.53. Additionally, the rail exit velocity is 69.2 ft/s.
With the motor selected and the materials decided, the propulsion system (housing) was designed. The bulkhead had a 5.38” diameter, 0.25” thick aluminum plate placed in front of a
3.1” outer diameter inner tube to accommodate the L850W motor. There will be two centering rings located along the inner tube with a 3.105” inner diameter and a 5.38” outer diameter.
These rings will be 0.25” thick 6061-T6 Aluminum. The thrust plate had the same dimensions as the centering rings and was located 0.25” from the end of the inner tube to allow for a retention system to be attached to the rocket. An exploded view of the motor mount can be found in Figure
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7. Figure 8 shows labeling of the components for the propulsion system. For a dimensional drawing, Figure 9.
Figure 7 - Exploded View of the Motor Mount
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Bulkhead Centering Rings Thrust Plate
Inner Tube Motor and Case
Figure 8 - Propulsion Components Labeled
Figure 9 - Dimensional Drawing for the Motor Mount
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The motor mount will be 21.25” long, including the bulkhead. This will leave enough room for the recovery system to reach the desired pressure needed for the system. The total estimated weight of the propulsion system with propellant is 10.421 lbf and the weight with no propellant is 5.861 lbf. All motor mount drawings can be found in Appendix A.
Recovery
The launch vehicle will utilize a dual-deployment recovery system with redundant altimeters to ensure that the vehicle lands safely at a reasonable distance from the launch site. A coupling tube will house the recovery electronic systems and serve to unite the two carbon-fiber body tubes. At apogee, a black powder ejection charge will pressurize the volume above the coupling tube, separating the rocket into two sections and deploying a ripstop nylon drogue parachute.
When the rocket has descended to an altitude of 1000 feet, a second black powder ejection charge will pressurize the volume below the coupling tube, separating the rocket again and deploying the main parachute, which will also be made from ripstop nylon. All three sections of the rocket will be tethered together using tubular nylon cord, which shall be protected from the ejection charges by flameproof fabric and attached to aluminum bulkheads using U-bolts.
At the heart of the recovery system are two PerfectFlite StratoLogger CF altimeters, shown in Figure 10. This particular model was chosen for its simplicity and cost-effectiveness; while the StratoLogger CF has a relatively limited set of functions, the alternatives considered were generally much more expensive and provided unnecessary features for the purposes of a simple dual-deployment operation.
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0.84”
2”
Figure 10 - PerfectFlite Stratologger CF Altimeter The altimeters will be powered independently of each other using two 9-volt batteries, and armed independently using two rotary locking switches accessible externally via two small holes in the airframe. These holes also serve to expose the altimeters to the external air pressure to allow accurate determination of the launch vehicle’s altitude. To preserve the redundancy of the system, each altimeter will operate on completely separate circuits, including separate igniters for each altimeter. Lead wires will connect the altimeter outputs to terminal blocks mounted to the outside of the coupler bulkhead. The terminal blocks allow for quick replacement of igniter wires. A block diagram showing the redundant recovery electrical system is shown in Figure 11.
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Figure 11 - Block diagram of major recovery system electrical components The altimeters, batteries, and arming switches will be mounted to a plywood sled inside the recovery bay. This plywood sled will be located on two threaded rods that are secured at each end to aluminum bulkheads, as shown in Figure 12. The bulkheads will mount flush to the coupling tube to isolate the altimeters from the pressure bursts associated with the black powder ejection charges. 5/16” steel U-bolts with steel backing plates will serve as attachment points for the 1” tubular nylon recovery harness.
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Figure 12 - Recovery bay bulkheads and hardware A very important consideration in the development of a recovery system for a high-powered rocket is the parachute configuration. The launch vehicle will utilize a system that houses the drogue and main parachutes in separate compartments on opposite sides of the recovery bay, as shown in Figure 13 and Figure 14. These compartments are bounded by aluminum bulkheads that are epoxied to the body tube and have identical U-bolts that serve as mounting points for the recovery harness. This configuration keeps all three sections of the rocket (nose, recovery bay, booster) tethered together after parachute deployment via the tubular nylon recovery harness, as shown in Figure 13. As per the decision matrices in the Component Alternatives section, a 36”
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Fruity Chutes Classic Elliptical parachute will serve as the drogue parachute, and a 96” Fruity
Chutes Iris Ultra parachute will serve as the main parachute. A 25’ length of tubular nylon will be used for the drogue parachute tether, and a 35’ length will be used for the main parachute tether.
Figure 13 – Exploded View; Recovery System
9” 12” 14”
35”
Figure 14 - Recovery system layout within airframe
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25’ 35’
Figure 15 – Tethering of Rocket Sections
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Payload
Electronic Payload
The electronic payload is located in the nosecone of the rocket. It contains a Atlus Telemega, which will record and transmit all flight data and a battery. The entire payload will be water proof. The location of the payload with respect to the nosecone can be seen in Figure 16. An annotated exploded view can be seen in Figure 17.
Figure 16 - Electronic Payload within Nosecone
Figure 17 - Exploded View of Electronic Payload
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The assembly of the electronic payload with respect to the nosecone is shown in Figure
18.
Figure 18 - Exploded Electronic Payload View with Nosecone In Figure 19, the assembled payload can be seen and in Figure 20 the mounting studs are clearly shown.
Figure 19 - Top View, Assembled Electronic Payload
Figure 20 - Bottom View, Assembled Electronic Payload
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Fragile Material Payload
The main objective of the fragile material housing payload is to protect an unknown object(s) throughout the duration of the flight. To do this, many designs were brainstormed and down selected. One main idea developed was to have a supporting material within a cylinder to house the object and keep it in place within the rocket. The main alternative was a spring damper system to reduce the force of the rocket felt by the payload entirely. Both ideas were combined, resulting in the current design.
Project ACE’s design consists of two concentric cylinders, one with supplemental material inside to hold the fragile material in place. The entire system consists of two different springs in series and parallel meant to absorb both large and small vibratory impacts. Concentric cylinders within the rocket tube to allow payload oscillation. Figure 21 shows allof the components of the payload spring system in an exploded view.
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Figure 21 - Payload Exploded View a) Cylinder 2
b) Cylinder 1
c) 12 CR1-400 Wire Rope Isolators
d) Baseplates
e) Hardware used to assembly the system
f) Main springs
g) Coupling baseplate
h) Outer most baseplate
i) U bolt holding assembly together
Project ACE plans to use 12 CR1-400 Enidine wire rope isolators. These will allow oscillation of Cylinder 1 to reduce forces transmitted to the fragile material, small vibrations, and overall acceleration. The concentric cylinders and wire rope isolators can be seen in Figure 22.
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Figure 22 - Components of the Main Payload a) Cylinder 1
b) Cylinder 2
c) Wire Rope Isolators
In Figure 22, Cylinder 2 will be concentric with the rocket’s main body tube as well as
Cylinder 1. Cylinder 2 will be made of aluminum while cylinder 1 will be made of ABS plastic and will be 3D printed then sealed. “c” shows the location of 3 of the 12 wire rope isolators. A dimensioned drawing of Cylinder 2 can be found in Appendix A.
Cylinder 1, shown in Figure 23 is designed to have inside dimensions of 3.5” diameter and 9” long. The dimensioned drawing can be seen in Appendix A. The maximum envelope given to teams in the project requirements is 6” long, however, we designed the cylinder to have 3” extra
36 | P a g e of thread for the cap to screw down at variable lengths. The reason for this is that Cylinder 1 will contain support material (material to be determined through testing) and when the unknown object(s) are placed within the container, the support material will be displaced the same volume as the object(s). To be sure the support material firmly holds the object(s), the lid will screw to variable distances to compress the material and object(s) regardless of their size.
Figure 23 – Payload Inner Cylinder Attached to Cylinder 2 are 5 base springs - designed to absorb most of the large impact forces such as the initial takeoff, parachute deployment, and landing. Prior to completing the first mathematical model utilizing Simulink, these springs were not included. However, the forces induced on Cylinder 1 and thus the fragile material were too large so a series and parallel spring system was created by introducing the 5 base springs. These springs can be seen in Figure
24 and are labeled a.
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Figure 24 - Springs Attached to the Main Payload and Bulkhead Attachment Then entire system consisting of Cylinder 1, Cylinder 2, 12 wire rope isolators, and 5 base springs, oscillates within the body tube of the rocket and is mounted to the bulkhead separating the payload bay and the recovery bay. This bulkhead can be seen in Figure 24 is labeled “b”.
The walls of the payload bay, as well as the outside of Cylinder 2, will be lubricated to ensure smooth translation during oscillation with graphite powder. Again, the exploded view of the entire system can be seen in Figure 21.
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Testing on the payload will not only decide the support material but will also test the validity
of the math model’s ability to select springs. Testing will be performed on the entire spring and
concentric cylinder system with the matrix seen below in Table 14.
Table 14 - Testing Matrix for Fragile Material
Testing Materials Weight # To be Tested
Egg 1.75 oz 2
Glass Stir Rod .2 oz 1
Glass Sheet N/A N/A
Light Bulb 1.1 oz 3
Small Ceramic/Porcelain China N/A N/A
Contact Support Materials (within Cylinder Weight per cubic ft. Density Grain Size Liquid/Solid Viscosity 1) Aerogel N/A N/A N/A N/A N/A Packing Peanuts .2 lb N/A Varies Solid N/A Styrofoam Pellets .2 lb N/A Varies Solid N/A Non-newtonian Fluid N/A N/A N/A Both Varies High Density Foam (cubes/sheets) Varies .93 g/cm3 As needed Solid N/A Spray in High Density Foam (injection Varies 3 lb/ft3 N/A Solid N/A system)
Testing is a primary part of this section as it will not only give validation to the design but
will also show shortfalls and areas of interest going into the demonstration. Testing for the
payload as a system will be done with drop tests at various heights associated with desired
impulse forces. The three main phases of flight to be tested will be the impact force, the main
parachute deployment force, and the force caused by the motor. The two impulse forces, the
parachute, and impact, will be estimated and then tested with drop tests. Additionally, the team
will be modifying the Charpy Impact Tester to give desired impulses. The engine thrust will be
tested by selecting points of interest from the thrust curve given by the manufacturer and
mimicking those forces at that point in time again with a drop test. Each test will be repeated
with the top filler material choices from the material and testing object matrix found in Table 14. 39 | P a g e
The math model for the payload started as an analysis of the system in the form of free body diagrams. The system drawing can be seen below in Figure 25.
Figure 25 - System Drawing and Force Balance
Figure 25 is the system diagram for the entire rocket (M1), the concentric cylinder and spring assembly (M2), and Cylinder 1 including the unknown object(s) (M3). This system was then derived into free body diagrams and accompanying force equilibrium equations seen below in
Figure 26- Figure 28.
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Figure 26 - Free Body Diagram for the Entire Rocket and Force Balance
Figure 27 - Free Body Diagram for the Spring Assembly and Force Balance
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Figure 28 - Free Body Diagram for Cylinder 1 and the Fragile Material and Force Balance The first round of Simulink models did not work since the model was under constrained.
For this reason, the team redesigned the Simulink model as a base excitation vibration model.
The main change this induced was that the only input in Simulink for M1 was the external force for the given situation. Three different models were made simulating three force events. A table showing these values is given below in Table 15.
Table 15 - Force Events for the Simulink Model
Force Input Thrust Thrust curve Main Parachute Deployment 400 ft. lb Landing 75 ft. lb
The way the math model helped us to select the needed springs was by selecting one of the inputs from Table 15 above and then iteratively selecting springs until one was found that fit our
42 | P a g e application. Different k (spring constant) and c (damping coefficient) values were inserted in the model for k1 and 2 and c1 and 2. This model can be seen below in Figure 29.
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Figure 29 – Simulink Mathematical Model 44 | P a g e
The mathematical model shows the Simulink model used to select springs for the system.
The approach taken was entering in estimated or known forces as the input for the base excitation model for mass 1 and then three graphical outputs were created, position, velocity, and acceleration. The acceleration graphs were then used to determine the overall force on the payload and springs were optimized by plugging in various k and c values to determine the best reduction in force and acceleration on the payload or mass 3. To determine the best spring selection, it was decided to perform iterations using said k and c values since those could be easily selected from standard commercial parts and then the system could be solved to find the resultant forces and accelerations desired. We decided the best spring was selected when the maximum displacement of the spring was reached without bottoming out and the smallest force and accelerations were transmitted to the payload.
The damping coefficients (c) present in the model were calculated from the manufacturer specifications that stated the damping was 5 percent of the spring constant. The final values for the constants as listed in Table 16.
Table 16 - Final Values for Constants
Final values (N/m)
kv 15761.4
cv 7.6
ks 4623.384
cs 4.121
Mission Performance Predictions
The main source of flight simulator data used for flight predictions was OpenRocket. This software’s flight simulation is based off of an atmospheric model that estimates variable
45 | P a g e conditions with changing altitude. This model assumes ideal gas for the air. This model also considered a wind model, importing the Kaimal spectrum equation and the assumption that the wind speed is uniaxial. Another assumption the program makes is that the earth is flat, which negates Coriolis effects. Additionally, turbulence intensities are based on wind farm load design standards, which may or may not translate to higher altitudes. With these models taken into consideration, the program runs a 4th order Runge-Kutta integration method with the following steps:
1. Initialize the rocket in a known position and orientation when time is equal to zero
2. Compute the local wind velocity and other atmospheric conditions
3. Compute the current wind speeds, angle of attack, and other flight parameters
4. Compute the aerodynamic forces and moments on the rocket
5. Compute the motor thrust and center of gravity
6. Compute the mass and moment of inertia of the rocket from linear and rotational
acceleration of the rocket
7. Numerically integrate acceleration to the rocket’s position and orientation during the time
step ∆t and update the time. (Niskanen, 2009)
The program computes steps 2-6 until the rocket has reached its end time which is
normally reaching the ground (Niskanen, 2009). This open source software is similar to
commercially available software such as Rocksim. OpenRocket originated at Helsinki
University of Technology as a Master’s Thesis by Sampo Niskanen (Niskanen, 2009).
Experiments working to prove that OpenRocket is accurate found that during one test on a B
size motor that the program over estimated the altitude by about 16%, and for a C size motor
altitude was over estimated by 7%. For another experiment, a larger motor was used and the 46 | P a g e
program under estimated altitude by 16%. However, the program was also compared to
commercially available software and it was found to be as accurate as Rocksim. In the same
experiment, Rocksim’s uncertainty was B motor – 24%, C motor- 19% and Larger motor-
12% (Niskanen, 2009).
For Project ACE’s rocket, the plan is to add around 50% of the allowable ballast to lower
the projected altitude to exactly 5,280 ft. Figure 30 shows the predicted altitude results from
OpenRocket. The inputs for the OpenRocket simulation can be found in Appendix B.
6000
5000
4000
3000 Altitude(ft) 2000
1000
0 0 20 40 60 80 100 120 140 160 180 Time (s)
Figure 30 - Predicted Altitude from OpenRocket Simulation The predicted altitude from the OpenRocket software is 5,379 ft. The inputs for the simulation were 4 mph for the average wind speed with a standard deviation of 0.4 mph. The inputs for the OpenRocket simulation can be found in Figure 31.
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Figure 31 - OpenRocket Flight Simulation Inputs The predicted altitude from the flight simulation of OpenRocket was compared to the predicted altitude using the same inputs and rocket design in Rocksim. This was to show that the altitude was predicted in multiple ways. The altitude that was predicted for the Rocksim model was 5,368 ft which was close to the predicted altitude from OpenRocket. Figure 32 for altitude predictions from the Rocksim simulation. This altitude showed a percent difference of 0.20% between the Rocksim simulation and the OpenRocket simulation. The OpenRocket value is used as the base because OpenRocket is the original program used to calculate the altitude of the rocket. Inputs for the Rocksim simulation are provided in Figure 33. Equation 1 showed how the percent difference was calculated and Equation 2 showed the calculated values for the percent difference.
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푂푝푒푛푅표푐푘푒푡 푉푎푙푢푒−푅표푐푘푠푖푚 푉푎푙푢푒 % 퐷푖푓푓푒푟푒푛푐푒 = Eq. 1 푂푝푒푛푅표푐푘푒푡 푉푎푙푢푒
5379푓푡−5368푓푡 % 퐷푖푓푓푒푟푒푛푐푒 = 푥 100 = 0.20% Eq. 2 5379푓푡
6000
5000
4000
3000 Altitude(ft) 2000
1000
0 0 20 40 60 80 100 120 140 160 180 Time (s)
Figure 32 - Predicted Altitude from Rocksim Simulation
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Figure 33 - Inputs for Rocksim Simulation The thrust curves produced by the simulations show the same thrust for the L850W motor. The thrust curve produced by Aerotech is shown in Figure 7. The thrust curve from the
OpenRocket simulation can be found in Figure 35 and the thrust produced in Rocksim simulation can be found in Figure 36. The components that were used in the simulations can be found in
Appendix B, along with weights of each component.
Figure 34 - Thrust Curve from AeroTech Motor 50 | P a g e
1400
1200
1000
800
600 Thrust Thrust (N)
400
200
0 0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5 Time (s)
Figure 35 - Thrust Curve for the L850W Motor in OpenRocket
1400
1200
1000
800
600 Thrust Thrust (N) 400
200
0 0 1 2 3 4 5 6
-200 Time (s)
Figure 36 - Thrust Curve for the L850W Motor in Rocksim
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The Center of Gravity (CG) is 69.92 in. from the tip of the nosecone. The Center of Pressure
(CP) is 90.46 in. from the tip of the nosecone. This produces a stability of 3.69 calipers. This was determined using OpenRocket.
CG CP
Figure 37 - Center of pressure and gravity Using the average atmospheric and weather conditions for an April day in Huntsville,
Alabama, an OpenRocket simulation was conducted to predict the performance of the recovery system. The drogue parachute provides a safe initial descent rate of 50.2 ft/s, which is suitable for keeping the landing site within walking distance of the launch site while also ensuring that the main parachute does not open under excessive speed. The rocket will impact the ground with a speed of 14.8 ft/s, giving each section of the rocket a kinetic energy under the maximum allowable 75 ft-lbf as shown in Table 17 below.
Table 17 - Kinetic energy of each section upon landing
Section Mass (lb) Kinetic Energy (ft-lbf) Nose Cone & 9.19 20.9 Payload Recovery Bay 4.32 12.66
Booster 10.03 29.4
Apart from these average atmospheric conditions, drift distances were simulated in
OpenRocket for different wind speeds as shown in Table 18. These distances assume a perfectly
52 | P a g e vertical launch angle with medium atmospheric turbulence. As the simulated drift distance for 20 mph winds is over the allowable distance of 2500 ft, the main parachute deployment altitude will be lowered using the altimeter’s built-in adjustment features in the event of excessive wind speeds on launch day.
Table 18 - Landing site distance from launch site by wind speed
Wind Speed (mph) Lateral Distance (ft)
0 7
5 576
10 1296
15 2087
20 3046
Safety
Overview
The University of Evansville, in conjunction with Project ACE and all team members, is dedicated to a successful launch, and, most importantly, safe operation of the rocket throughout all phases of the project. Led by Safety Officer, Bryan Bauer, the team members will be saturated with information regarding proper safety protocols for each stage of the project. In addition to this, all team members will be briefed on the hazards that are specific to the materials they will come in direct contact with so that accidents and injury can be prevented. Furthermore, material data sheets (MSDS) will be available to all students in the working area, so that potential hazards can be identified before construction begins.
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During the construction and fabrication phase of the project, students will work in groups of no less than two, to ensure that at least one team member would be able to provide immediate assistance and call for help in the event that an accident occurs. Additionally, the team safety officer will monitor use of personal protection equipment (PPE), such as glasses and gloves amongst other things, during construction to ensure all team member are safe. The team safety officer will also ensure that the energy systems lab is equipped with working smoke detectors and fire extinguishers as well as first aid kits.
During the sub-scale and full-scale testing of the rocket, all team members will wear safety glasses and will maintain a safe distance from the launch pad. Due to the risks associated with various facets of the rocket, checklists will be developed and reviewed before final assembly and launch to guarantee safety of all team members and spectators. Additionally, the team will work together to construct a hazard analysis which will be used to identify risks, their causes, and proposed mitigations in order to minimize the chance of accident and injury, and ensure safe operation. This focus on safety and education of all team members will create optimal working conditions, which ultimately will keep the project on schedule and allow for safe and successful launch.
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Final Assembly Checklist
Initials Check-Off Points
______Check rocket tube for cracks, bumps, abrasions or any other imperfections
that could have been acquired during construction or transport that could
adversely affect the flight of the rocket. ______Check parachute for any inadequacies or tears that could alter deployment
and safe landing.
______Ensure that the parachute is packaged properly inside the rocket tube.
______Check payload for any cracks or chips that could have been acquired
during transport.
______Check motor and casing to ensure it is not wet or containing any visible
imperfections that would cause a misfire or deviation from the ideal flight
path. ______Ensure recovery harness is properly attached for flight readiness. ______Check motor mount for structural integrity. ______Check primary fins for cracking or bowing. ______Check thrust plate and couplers for solid attachment and structural
integrity to ensure proper flight.
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______Check avionics bay for proper functioning to ensure noting was broken or
altered during transport
______Check nosecone for structural integrity and secure attachment to the rest
of the rocket.
______Insert motor into casing and check for secure fit
______Ensure all connections of the rocket are solid and cohesive
UE SLI Safety Officer Signature
______
UE SLI Team Lead Signature
______
UE SLI Adult Educator Signature
______
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Launch Procedures Checklist
Launch Procedures Checklist
______Ensure a safe working area before unloading the rocket and bringing it to
the launch pad.
Check the safety and readiness of team members by ensuring all team ______members have on safety glasses and other proper PPE for the part of the
rocket they will be handling
Visually inspect the rocket for proper connections between all sections ______before placing on the launch pad. ______Test electronics (i.e. camera, altimeter, etc.) to ensure they are fully
functional and turned on before launch
______Check launch pad and guide rails for readiness
______Place rocket on launch pad
______Have non-essential team members move away from the launch pad to the
safe viewing distance
______Arm the rocket motor for ignition
______Disarm all safeties on the rocket
Have remaining team members move to safe viewing distance to watch ______the launch
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______Check with Range Safety Officer (RSO) to ensure all codes and rules are
met and the rocket is clear for launch.
______Initiate rocket ignition.
Check for proper ignition ______
UE SLI Safety Officer Signature
______
UE SLI Team Lead Signature
______
UE SLI Adult Educator Signature
______
*Note: The launch procedures checklist will be edited during the course of the project to include more detail as the team learns more about standard launch procedures and the setup of the rocket.
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Personnel Hazard Analysis
A preliminary personnel hazard analysis was conducted to identify hazards, causes and
resulting effects. This analysis was created make team members aware of potential hazards, and
lists mitigations to reduce the chance of risk or injury during the course of the project. This
analysis is summarized in Table 21.
Table 19 - Personnel Hazard Analysis
Risk/Hazard Effect/Severity Severity Likelihood Mitigation and Control Inhalation of toxic fumes, accidental Work in well ventilated Epoxy ingestion, or contact with skin leading Minor High spaces to irritation or rash
Inhalation of dust particles from Wear mask when sanding Dust sanding or machining operations Minor High to avoid inhaling dust Particles resulting in breathing problems particles
Heavy Tools Improper handling of shop tools or Ensure proper training for and machining operations leading to all team members working Significant Medium Machinery personal injury or destruction of with any tool or machinery in Lab equipment in shop
Properly transport motor Rocket Exposure to rocket fuel in contact with Major Medium from offsite location to Propellant skin leading to irritation and burns launch site Gases may be toxic if exposed in areas Store in portable fireproof Black with inadequate ventilation. Also keep Major Low case to keep away from Powder away from open flame, sparks, and heat fire and high temperatures
Ensure at least one teammate is working Craft and Cuts leading to injury as a result of alongside the person doing Exacto precision cutting operations on fins or Minor Medium the cutting. Practice safe Knives other pieces of the rocket body cutting procedures by cutting away from body.
59 | P a g e
Store a fire extinguisher in Burns, significant and/or fatal injury, or the room where the rocket damage to school from fire as a result will be constructed. If an Fire Major Low of faulty wiring, or improper handling object starts to overheat, let of the motor and black powder it cool and have the fire extinguisher ready Be aware of surroundings Bruises, cuts or scrapes from when operating the Handheld mishandling of basic handheld shop Significant High handheld tools and ensure Tools tools such as hammer or saw proper training before any construction is undertaken.
Failure Modes and Effects Analysis
A preliminary Failure Modes and Effects Analysis of the proposed design of the rocket,
payload, payload integration, launch support equipment, and launch operations, which can be
seen in Table 22, was completed to identify hazards, effects and proposed mitigations.
Table 20 - Failure Modes and Effects Analysis
Proposed Risk/Hazard Effect Severity Likelihood Mitigation Properly transport motor from offsite Motor Improper handling or storage of location to launch Handling/Accidental motor resulting in accidental or Major Low site. Ensure proper Ignition unexpected ignition connections before launch Maintain safe distance from launch pad. Have Failure of motor to ignite and Launch Failure Significant Low team mentor/safety launch rocket properly officer inspect rocket on launch pad Failure of the secondary parachute Maintain safe Main Parachute to deploy leading to freefall or Major Low distance from Deployment Failure unstable flight of rocket back to launch pad the ground Failure of the initial parachute to Maintain safe Drogue Parachute deploy leading to freefall or Significant Low distance from Deployment Failure unstable flight of rocket back to launch pad the ground
60 | P a g e
Failure of the rocket to maintain Maintain safe Instability During its projected flight path due to Major Low distance from Flight unforeseen design flaw or in flight launch pad malfunction Verify all Altimeter or Other electronics work Electronics in Potential short circuiting or harm properly before Avionics Bay Minor Medium to spectators below launch and are Malfunction/Fall firmly attached to Off the rocket Run multiple tests to ensure proper Coupler Excessively Failure of parachute to deploy amounts of black Major Low Tight leading to damage to rocket powder is used to allow rocket to separate Take caution when inputting payload into rocket before Payload Not Inability to return materials launch and ensure Minor Medium Secured Properly without breaking all items are properly sealed and secured before launch
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. This analysis is shown below in Table 23.
Table 21 - Environmental Consideration Analysis
Risk/Hazard Effect and Impact Severity Likelihood Mitigation and Control Vehicle Effects on Environment
When epoxying various pieces of the rocket Work in well ventilated Epoxy Fumes together, harmful fumes Minor High spaces and dispose of are released into the waste properly atmosphere
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Wear mask when sanding Small dust particles from to avoid inhaling dust sanding or machining particles and try to operations are released into Dust Particles Minor High contain dust when the environment which can sanding opposed to freely result in breathing releasing it into problems surrounding air. Upon ignition, the motor Place flame resistant reaches high temperatures material beneath the Rocket Motor and hot exhaust is released, launch pad to avoid Major Low Ignition which could potentially burning the immediate burn the areas where the surroundings or starting a rocket is launched or lands fire
If various pieces of the rocket do not stay intact Ensure fully functioning during decent, or the parachutes before launch parachutes do not operate via pre-launch checklist properly, pieces of the and check that all Debris from Rocket Significant Low rocket could break off components of the rocket during flight or upon and payload are impact and be irretrievable, accounted for upon leading to minor return. environmental harm.
Environmental Effects on Vehicle
Precipitation and moisture within the rocket could Avoid launching rocket affect the structural in wet conditions and Water integrity of the rocket, or Significant Low ensure a dry area for could lead to malfunctions storage and transport of the electronics housed in the avionics bay
Strong wind or unpredictable wind gusts Avoid launching rocket can cause the rocket to on days where high speed Wind deviate from its ideal flight Significant Medium winds or unpredictable, path and can lead to strong wind gusts are damage to the rocket and present potential harm to spectators
62 | P a g e
Humidity can lead to moisture in the body of the rocket which can lead to Store rocket in a dry area corrosion and weakening to avoid moisture Humidity of various materials used to Minor Low entering the rocket over construct the rocket. It can time via humid air also negatively impact on- board electronics
General Risk Assessment
Finally, a general risk assessment 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. Seen in Table 24.
Table 22 - General Risks Associated with the Project
Risk/Hazard Effect Impact Likelihood Proposed Mitigation Value
Due to the new nature of the The team will work with project to this team specifically, if faculty members as well the team is unable to find as local rocketry club Limited valuable insight from external High Medium members in order to gain Resources sources, the design and a better understanding of performance of the rocket could rocketry and develop a suffer functional rocket.
The team and its adult Lack of flexibility in the budget educators will apply for could lead to the team being grants and fundraise to Tight or forced to use parts that are not provide the team with a High High Minimal Budget optimal, or being unable to flexible budget beyond replace parts of the rocket that are the normal amount of broken during testing money allotted to the project by the school
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Inability of the team to keep up Team members will fill with the initial schedule set forth out weekly time cards Mismanagement in the task breakdown could lead and log their hours in the Medium Low of Time to major delays, poor quality of task breakdown in order work, or the rocket not being to ensure everyone completed by competition remains on schedule
Failure to properly account for the work needed to complete the There will be constant Underestimation project could lead to the project communication amongst of Scope of running behind schedule and Medium Low all team members and Work various facets of the rocket not with NASA to ensure the being completed in a quality scope of work is clear manner The team will design and downselect with safety in mind, and will clearly identify all safety Increase in Adding material to the rocket in measures before Safety order to increase safety will result Low Medium construction so that Regulations in an increase in expenses additional, last-minute safety measures do not have to be taken that will inflate the budget.
Project Plan
Requirements Compliance
Table 23 - Requirement Compliance
NASA Requirements Handbook Summarized Verification Description of Verification Plan Number Requirement Method(s) The vehicle shall deliver the science or The rocket team will utilize OpenRocket, engineering payload Test 1.1 RockSim, CFD, & test flight data to to an apogee altitude Analysis achieve an accurate prediction of altitude. of 5,280 feet above ground level (AGL).
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NASA Requirements Handbook Summarized Verification Description of Verification Plan Number Requirement Method(s) The vehicle shall carry one commercially available, barometric The rocket will house a Atlus Metrum altimeter for recording TeleMega altimeter in the nosecone to 1.2 Inspection the official altitude record the official altitude used in used in determining determining the altitude award winner. the altitude award winner. All recovery electronics shall be Batteries & altimeter will be purchased 1.3 powered by Inspection from online rocketry sources. commercially available batteries. The launch vehicle shall be designed to be recoverable and The rocket is reusable in design because reusable. Reusable is Test our team is using a motor that has refuels 1.4 defined as being able Inspection that can be reloaded into the motor under to launch again on the supervision. same day without repairs or modifications. The launch vehicle The launch vehicle will have 3 shall have a maximum independent sections: the aft body tube, 1.5 Inspection of four (4) the bow body tube and nosecone, and the independent sections. coupler. The launch vehicle Inspection 1.6 shall be limited to a The launch vehicle shall be a single stage. Demonstration single stage. The launch vehicle The launch vehicle will be designed with shall be capable of Inspection an efficient and quick to construct design 1.7 being prepared for Demonstration that requires fewer than 4 hours to flight at the launch prepare. site within 4 hours.
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NASA Requirements Handbook Summarized Verification Description of Verification Plan Number Requirement Method(s) The launch vehicle shall be capable of remaining in launch- ready configuration at The launch vehicle design will ensure all the pad for a 1.8 Test components have a life of greater than 1 minimum of 1 hour hour without loss of functionality. without losing the functionality of any critical on-board component. The launch vehicle shall be capable of Inspection The ignition system will be using a 12 1.9 being launched by a Test volt direct current firing system. standard 12-volt direct current firing system. The launch vehicle shall require no There will be no external circuity for the external circuitry or ignition system because it will be a special ground support ground based ignition system being 1.10 Inspection equipment to initiate placed underneath the rocket before launch (other than launch with 300 ft of cord between the what is provided by igniter and the controller. Range Services).
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NASA Requirements Handbook Summarized Verification Description of Verification Plan Number Requirement Method(s) The launch vehicle shall use a commercially available solid motor propulsion system using ammonium perchlorate composite propellant (APCP) Inspection The motor being used is a solid fuel 1.11 which is approved and motor from AeroTech. The motor is the
certified by the L850W. National Association of Rocketry (NAR), Tripoli Rocketry Association (TRA), and/or the Canadian Association of Rocketry (CAR). Pressure vessels on 1.12 the vehicle shall be Inspection No pressure vessels will be used. approved by the RSO. The total impulse provided by a The motor will produce an impulse of University launch Test 1.13 3695 N-s which is below the specified vehicle shall not Analysis total impulse that is allowed. exceed 5,120 Newton- seconds (L-class). The launch vehicle shall have a minimum Test The launch vehicle will have a static 1.14 static stability margin Analysis stability margin of 2.67. of 2.0 at the point of rail exit. The rocket team will utilize OpenRocket, The launch vehicle RockSim, CFD, & test flight data to shall accelerate to a Test 1.15 achieve an accurate prediction of minimum velocity of Analysis minimum velocity at rail exit. The current 52 fps at rail exit. value is 67.2 ft/s.
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NASA Requirements Handbook Summarized Verification Description of Verification Plan Number Requirement Method(s) All teams shall successfully launch Test A subscale model with comparable 1.16 and recover a subscale weights, lengths, and masses will be Analysis model of their rocket launched prior to the CDR. prior to CDR. All teams shall successfully launch The project schedule will ensure a full- and recover their full- Test 1.17 scale rocket launch occurs before the scale rocket prior to Analysis FRR. FRR in its final flight con- figuration. The rocket will have 3 bolts holding the Any structural nosecone to the bow body tube and shear protuberance on the Test pins holding the coupler to the bow and 1.18 rocket shall be located Analysis aft body tubes. These structural aft of the burnout protuberances are all located aft of the center of gravity. burnout center of gravity Inspection The launch vehicle will follow all 1.19 Vehicle Prohibitions Test prohibitions laid out in section 1.19 of the Analysis 2017 SL NASA Student Handbook. Vehicle must deploy a drogue parachute at Demonstration Dual-deployment altimeters will be 2.1 apogee, followed by a programmed to fire ejection charges at Inspection main parachute at a apogee and at ~1000 feet. much lower altitude. A successful ground ejection test for both parachutes must be Multiple ejection tests will be conducted 2.2 Test conducted prior to prior to sub- and full-scale launches. sub- and full-scale launches.
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NASA Requirements Handbook Summarized Verification Description of Verification Plan Number Requirement Method(s) No part of the launch vehicle may have a Analysis Parachute sizes will be optimized to 2.3 kinetic energy of minimize kinetic energy at ground Demonstration greater than 75 ft-lbf impact. at landing. Recovery electrical circuits must be Recovery electronics will be located in a 2.4 Inspection independent of separate, shielded coupler. payload circuits. Recovery system must include redundant, Two PerfectFlite Stratologger CF 2.5 Inspection commercial altimeters will be used. altimeters. Motor ejection cannot be used for primary or Demonstration Black powder ejection charges will be 2.6 secondary Inspection used to eject parachutes. deployment. Each altimeter must be armed by a Locking rotary switches and LED 2.7 dedicated switch Inspection indicators will be used to confirm the accessible from the state of the recovery electronics. rocket exterior. Each altimeter must Separate 9-Volt batteries will be used to 2.8 have a dedicated Inspection power the altimeters. power supply. Each arming switch Locking rotary switches will be used to 2.9 must be lockable to Inspection arm each altimeter. the “ON” position. Removable shear pins must be used to seal Threaded nylon shear pins will be used to 2.10 Inspection the parachute seal the parachute compartments. compartments.
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NASA Requirements Handbook Summarized Verification Description of Verification Plan Number Requirement Method(s) Tracking device(s) Test All parts of the launch vehicle will be must transmit the tethered together; position will be 2.11 position of any parts Demonstration transmitted via a flight computer in the of the launch vehicle Inspection nosecone. to a ground receiver. Recovery system electronics must not Test Recovery electronics will be located in a 2.12 be adversely affected Inspection separate, shielded coupler. by any other on-board electronics. Design container Math model is used to develop spring capable of protecting system in conjunction with a concentric 3.4.1 an unknown object of Testing cylinder model to provide sufficient unknown size and vibration dampening and force reduction. shape. The spring and concentric cylinder design will be tested with a matrix of different Object must survive support materials as well as testing 3.4.1.2 Testing duration of flight materials to assure the unknown object(s) can survive the flight during demonstration. Once the object is obtained, it must be Support material within cylinder 1 that sealed in its housing allows object to be inserted and not spill 3.4.1.4 until after the launch Demonstration any material such as a high viscosity fluid and no excess material or malleable solid. may be added after receiving the object. Final assembly and pre-launch checklists Each team shall use a will be created and reviewed at the 4.1 launch and safety Inspection appropriate time to ensure safe launch of checklist the rocket and all members involved in the launch
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NASA Requirements Handbook Summarized Verification Description of Verification Plan Number Requirement Method(s) Each team shall identify a student safety officer who The team has appointed a safety officer to shall be responsible monitor the safety of the team throughout 4.2 for the safety of the Inspection the project and ensure all federal rules team and ensure all and laws are met. proper rules and guidelines are followed The team safety officer shall monitor team activities with an emphasis on safety The team safety officer will monitor the throughout the design, progress of the project emphasizing the 4.3 Inspection construction, and proper safety procedures for the current testing of the rocket stage of the project. by maintaining MSDS sheets and hazard analyses Each team shall appoint a mentor who has certification and is The team has assigned an school faculty in good standing with member to mentor the project to provide the NRA. This 4.4 Inspection valuable insight on the rocket design and member will be construction as well as assume full designated as the liability of the rocket. individual owner of the rocket and assumes liability During test flights, Team will converse with RSO at local teams shall abide by rocketry club to ensure all of their 4.5 the rules and guidance Demonstration chapter’s rules and regulations are abided of the local rocketry by. club's RSO
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NASA Requirements Handbook Summarized Verification Description of Verification Plan Number Requirement Method(s) Team will converse with NASA lead Teams shall abide by safety officer and thoroughly research all 4.6 all rules set forth by Demonstration rules and regulations set forth by the FAA the FAA to ensure all rules and regulations are abided by. Students shall do The team will continuously demonstrate 100% of the project Demonstration an independently managed and executed 5.1 excluding motor / Inspection project. The team lead will routinely black powder monitor this quality. handling. Documents for scheduling, budget A detailed project plan 5.2 Demonstration tracking, outreach, and safety will be shall be maintained. continuously updated and reported. Foreign National The team lead will ensure that any 5.3 members shall be Inspection Foreign National members are clearly identified by the PDR. indicated in the PDR. All team members It will be checked that a list of team attending launch week members, with indications of those 5.4 Inspection shall be identified by attending launch week, will be included the CDR. in the CDR. The educational The Educational Engagement lead shall engagement confirm that all documentation has been 5.5 Inspection requirement shall be received and approved by NASA prior to met by the FRR. the FRR. The team shall Team members will periodically confirm develop and host a 5.6 Test that the website is functioning as intended website for by opening each posted document. documentation. The team shall post & make available for The team lead shall confirm that all 5.7 download all Inspection documents are posted prior to the deliverables by the specified date. specified date. All deliverables must The team lead shall confirm that all 5.8 Inspection be in PDF format. documents posted are in PDF format.
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NASA Requirements Handbook Summarized Verification Description of Verification Plan Number Requirement Method(s) A table of contents The team lead shall ensure that a table of 5.9 must be included in all Inspection contents is located at the start of each reports. report. Page numbers shall be Page numbers shall be checked to the 5.10 provided in each Test table of contents to ensure continuity report. throughout the report. The team shall provide Demonstration Videoconference rooms will be reserved 5.11 videoconference and trialed immediately prior to each Test equipment needed for design review. reviews. All teams shall use The team shall design the rocket to utilize 5.12 launch pads provided Demonstration 1515 12’ launch rail. by the SLS provider. If software or applications are created The team must (not planned) the team will abide by 36 implement the EIT 5.13 Demonstration CFR Part 1194. Otherwise, all accessibility components containing software will be standards. checked to ensure compliance.
Team requirements have been developed in addition to the NASA requirements. These can be seen in Table 24.
Table 24 - Team Requirements Team Requirements Verification Number Requirement Description of Verification Method Method All reports shall be Reports shall be completed, according to compiled at least three team schedule, prior to NASA due dates 1 Demonstration days prior to NASA to allow for revision time and mitigate due dates. risk of late submissions.
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Team Requirements Verification Number Requirement Description of Verification Method Method At each team meeting, every sub-section Each member of the lead will review the status of their section team shall have a 2 Inspection with the entire team. The team leader working knowledge of will confirm that the information each subsystem. presented is sufficient. Safety shall be made The safety officer will periodically ask 3 the team’s first Test team members what the most important priority. aspect of the project is. All altimeters shall be flown on sub-scale Altimeters shall be in and full scale flight tests. Altitude 4 Test good working order. readings will be compared to confirm consistency. The tracking system shall be flown on The tracking system the sub-scale and full scale flight tests. 5 shall be in good Test This will be used to find the rockets thus working order. confirming its operation. A solid output signal Test All altimeters will be triggered while 6 must be given from voltage is read on the output. This output Analysis triggered altimeters. will be read to confirm it is acceptable. All circuits shall be All circuits will be confirmed at each 7 Demonstration checked prior to use. node to ensure connections. The main parachute shall have an Impulse for the apparatus (strain gauge) attached to it parachute deployment Test that enables a force to be read as it opens 8 shall be determined Analysis at high speed. This will cut down in the experimentally. large ambiguity that exists in estimating an impulse value. The spring constant shall be determined A spring constant for using forces related to what is parachute cords shall Test experienced with parachute opening. 9 be determined Analysis This helps when estimating energy experimentally. absorption by the cord when the chute opens. Payload must reduce From the mathematical model, 10 force felt by object(s) Testing appropriate springs will be selected to by 50 % induce oscillation and reduce force.
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Team Requirements Verification Number Requirement Description of Verification Method Method From the mathematical model, Payload must reduce appropriate springs will be selected to 11 acceleration of Testing insure acceleration graphs show 35 object(s) by 35 % percent reduction from inputs.
Budget
The budget was able to be based on a detailed parts list due to much preliminary work by the
Project ACE team. This list can be seen in Appendix C. To create the budget, the team first broke down the rocket into a number of sections (i.e. recovery, aerodynamics, etc.) Then the aforementioned parts list was created for each section. The total cost of each section then had a contingency budget implemented based on the risk of that section. The aerodynamic section can be taken as an example. The parts list calls for $1,051.67 in components. $348.33 was added to this amount to mitigate component failure risk (a new nosecone can be quickly purchased if necessary, for example.) The sum of these for all sections of the rocket is shown in the
“Forecasted Amount” column of Table 25. Propulsion and travel were the only ‘major’ budgetary change from the proposal. Propulsion increased by nearly $1,000 due to unforeseen motor costs while travel costs decreased by nearly the same amount due to the University of
Evansville Department of Engineering agreeing to cover advisor (professor) travel costs.
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Table 25 - Section Level Budget
Item Forecasted Amount Percent of Total Operating $300.00 3% Travel / Lodging $2,730.00 26% Launch Pad $220.00 2% Aerodynamics (Body) $1,400.00 13% Propulsion $2,500.00 24% Main Payload $500.00 5% Electronic Payload $670.00 6% Recovery $1,150.00 11% Scale Model $1,050.00 10% Educational $100.00 1% Engagement Total $10,620.00 100%
Project ACE’s funding plan has had a slight re-allocation of funding since the proposal. Less funding will be received through the student government association and more funding will be received through the college of engineering. The breakdown of project funding is shown in
Table 26.
Table 26 - Funding Sources
Funding Amount Remaining NASA Grant $5,000.00 $5,620.00 SGA $2,730.00 $2,890.00 U.E. ENGR $2,890.00 - Total $10,620.00
Schedule
The team has broken up the project in numerous tasks. The full extent of these tasks and associated schedule can be found in Appendix D. To be concise, the team has combined many of these tasks into “activities” and developed a Gantt chart (Figure 38). For each of these
“activities”, Project ACE is currently on schedule or ahead of schedule. In the Gantt chart, the
76 | P a g e yellow column represents the current week. The vertical green line indicates where the team is at for each task. For example, the team is three weeks ahead of schedule for the Rocksim model.
Figure 38 - Gantt Chart
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In addition to the project tasks/activities the team has compiled a list of critical dates. These dates are crucial to the success of the project and are listed in Table 27.
Table 27 - 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 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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References
1. 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
2. Niskanen, S. (2009). Development of an Open Source model rocket simulation software.
OpenRocket. Helsinki: HELSINKI UNIVERSITY OF TECHNOLOGY.
3. Ring, C. (2016, 9 27). Launch Crue. Retrieved from LaunchCrue.org:
https://www.launchcrue.org/
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Appendix A – Machine Prints Dimensioned Drawings
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Appendix B – OpenRocket Simulation Inputs for OpenRocket Flight Simulation
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Appendix C – Parts List Parts List
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Section Item Description Part Number Manufacturer Lead Time (days) Quantity Price (ea) Price (total) Nose Cone 5.5" FIBERGLASS 4:1 OGIVE NOSE CONE 20540 Apogee 1 $ 84.95 $ 84.95 Body Tube 5.5" x 48" Carbon Fiber Airframe Wildman Rocketry 30 days 2 $ 350.00 $ 700.00 Fins G10 FIBERGLASS SHEET 1/4" X 1 SQ FT 14154 Apogee 4 $ 54.00 $ 216.00 Nose Cone Threads Adhesive Mount Nut 98007A013 McMaster 10 $1.44 $ 14.44 Nose Cone Bolts Stainless Steel Button-Head Socket Cap Screws 98164A134 McMaster 50 $ 0.13 $ 6.28 Rail Buttons LARGE AIRFOILED RAIL BUTTONS (FITS 1.5" RAIL - 1515) 13069 Apogee 3 $ 10.00 $ 30.00
$ - Aerodynamics $ -
$ 1,051.67 Motor AeroTech L850W 7525S AeroTech 1 $ 1,420.00 $ 1,420.00 Retaining System Aero Pack 75mm Retainer - L 24054 Apogee 1 $ 47.08 $ 47.08 Epoxy G5000 Rocketpoxy 2-pint package 30511 Apogee 2 $ 38.25 $ 76.50 Motor Mount 75mm Blue Tube 48" 10504 Apogee 1 $ 29.95 $ 29.95 Motor Reloads AeroTech L850W Refuels 12850P AeroTech 3 $ 199.99 $ 599.97
Propulsion Centering Rings and Bulkheads .250" Aluminum Plate 6061-T651 2x4 P314T6 Metals4uOnline 7 1 $ 181.50 $ 181.50
$ 2,355.00 5.5" Aluminum Bulkplate 25096 MadCow Rocketry 4 $ 25.00 $ 100.00 U-Bolts w/mounting plates for use with aluminum bulkhead (pack of 5) 3043T68 McMaster 1 $ 5.89 $ 5.89 Electronics bay coupler 5.5" OD, bulkheads, rails 10526 Apogee 1 $ 56.95 $ 56.95 Igniter terminal block for easy igniter replacement 9191 Apogee 2 $ 3.41 $ 6.82 Crimp Connector - Radioshack 2 $ 5.00 $ 10.00 Ejection well 2-pack PVC wells for black powder 3068 Apogee 2 $ 3.15 $ 6.30 Parachute Protector 18" Nomex flameproof cloth 29314 Apogee 2 $ 10.49 $ 20.98 Tubular Nylon Recovery Harness 30351 Onebadhawk 60 $ 1.10 $ 66.00 Shock Cord Protector 30" flameproof sheath 29300 Apogee 2 $ 12.95 $ 25.90 Rotary Switch lockable switch 9128 Apogee 2 $ 9.93 $ 19.86 Shear Pins Nylon, threaded (10 pack) 29615 Apogee 10 $ 3.10 $ 31.00
Recovery 0.3125" Quck Link Delta-shape link eyebolts, chutes, and cord - Giant Leap Rocketry 6 $ 11.54 $ 69.24 36" Drogue Chute 36" Classic Elliptical Chute 29165 Apogee 1 $ 95.17 $ 95.17 96" Main Chute Torroidal, 2.2Cd, Ripstop Nylon 29185 Apogee 1 $ 346.53 $ 346.53 Stratologger CF Main & Backup 9104 Apogee 2 $ 58.80 $ 117.60 Quest Q2G2 igniter 6-pack of igniters - Quest 4 $ 5.00 $ 20.00 Parachute Slider slows parachute deployment Giant Leap Rocketry 1 $ 13.22 $ 13.22 Black Powder - Gun Store 1 $ 20.00 $ 20.00 9 Volt Battery - Radioshack 4 $ 10.00 $ 40.00 22 Gague Wire - Radioshack 3 $ 1.00 $ 3.00 $ 1,074.46 Atlus Metrum TeleMega From csrocketry.com Atlus Metrum 21 1 $ 406.10 $ 406.10 Starter Pack From csrocketry.com Atlus Metrum 0 1 $ 100.00 $ 100.00 Arrow 440-3 Yagi Antenna get from link in start pack page Yagi 0 1 $ 50.00 $ 50.00 SMA to BNC adapter From csrocketry.com Atlus Metrum 0 1 $ 10.00 $ 10.00 Washers McMaster, For Spacing & Damping 90133A005 McMaster 3 1 $ 6.81 $ 6.81 O-Ring Bolts 10-24, 9/16in 91864A091 McMaster 3 1 $10.69 $ 10.69 Altimiter Bolts 5-40, 5/8in 91251A130 McMaster 3 1 $8.98 $ 8.98
ElectronicPayload Studs for Ballast .25 x 40, 1 in long 98750A011 McMaster 3 4 $1.07 $ 4.28
$ 596.86 Estimated Maximum $ - Exact Components TBD $ - Blue Tube (Testing) 5.5" x 48" Carbon Fiber Airframe 10506 Apogee - 1 $ 56.95 $ 56.95 Outer Cylinder (Coupler) 5.36" OD, 5.217" ID Blue Tube 13106 Apogee 1 $ 18.95 $ 18.95 Fastening Nuts For 3/8" x 16 Bolt, 1/4" Height 91813A190 McMaster 1 $ 11.08 $ 11.08 Fastening Bolts 3/8" x 16 x 1" 91251A621 McMaster 1 $ 8.62 $ 8.62 Base Washer 0.5" ID 1.25" OD 98026A114 McMaster 3 $ 7.47 $ 22.41 Studs 3/8" x 1" Length 95475A624 McMaster 1 $ 9.41 $ 9.41
Recovery Bolts 3/8" x 1.25" Length 91251A626 McMaster 1 $ 9.27 $ 9.27 Main Payload Main Recovery Nuts 3/8" Flanged 96282A103 McMaster 1 $ 6.98 $ 6.98 Spacing Pipe 5.25" OD and 4.75" OD 8486K954 McMaster 1 $ 57.46 $ 57.46 Springs Part Number 866, custom, century spring corp 5 $ 30.00 $ 150.00
$ 351.13
Educational Engagement Supplies TBA - - $ 100.00
Safety / Educational Engagement $ 100.00 RockSim Temporary, 1 Seat License 1123 Apogee 0 1 $ 20.00 $ 20.00 Shirts Notable Sponsors 3 $ 43.33 $ 130.00 Hotel (Group A) Apr. 5 - 8, 2/Room, Avg. $120/night 10 People - - 5 $ 360.00 $ 1,800.00 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 15 $ 10.00 $ 150.00
Administrative / Travel / Administrative $ 3,030.00 1515 Rail 1515 Extruded Al., 145" 16U252 Grainger 2 1 $ 140.71 $ 140.71 Rail Bracket 90 Degree 5 Hole Bracket 47065T271 McMaster 2 4 $ 9.74 $ 38.96 Bolts M10 x 20 x 1.5 91290A516 McMaster 2 1 $ 6.41 $ 6.41
Shipping (McMaster) $ 11.51 Launch Pad Launch $ 197.59 Body Tube 3" CARBON FIBER TUBING 60 INCHES LONG CFT3.0-60 Wildman 30 days 1 $ 218.50 $ 218.50 Nose Cone 3" FIBERGLASS 4:1 OGIVE NOSE CONE 20520 Apogee 1 $ 30.95 $ 30.95 Fins G-10 Fiberglass Sheet 0.125" (1/8") 12" x 24" Giant Leap 1 $ 52.49 $ 52.49 Rail Buttons LARGE AIRFOILED RAIL BUTTONS (FITS 1.5" RAIL - 1515) 13069 Apogee 2 $ 10.00 $ 20.00 Motor I435T 3836SC AeroTech 1 $ 149.99 $ 149.99 Motor Reload I435T Reloads zero94314 AeroTech 2 $ 54.99 $ 109.98 InnerTube 38mm BlueTube 10501 Apogee 1 $ 16.49 $ 16.49 Centering Rings/ Bulkhead Same as full scale/ use same sheet P314T6 Metal Depot - - 75mm Electronics Bay 10524 Apogee 1 $ 39.93 $ 39.93 ScaleModel 48" Main Chute 29167 Apogee 1 $ 126.85 $ 126.85 18" Drogue Chute 29162 Apogee 1 $ 56.90 $ 56.90
Subscale Shipping $ 38.95
Total Total $ 861.03
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Appendix D – Task Breakdown Task Breakdown
Project ACE Detailed Task Breakdown
Task* Responsible Start Date End Date 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 1.1.1 Create Standards for Proposal David May. 25, 2016 May. 25, 2016 Jun. 1, 2016 May. 5, 2016 1.1.2 Write Proposal David Sept. 1, 2016 Sept. 1, 2016 Sept. 27, 2016 Sept. 26, 2017 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 1.2.2 Write Preliminary Design Review David Oct. 5, 2016 Oct. 5, 2016 Oct. 26, 2016 1.3 Critical Design Review (Report) David - - - 1.3.1 Create Standards for Critical Design Review David Oct. 28, 2016 Nov. 2, 2016 1.3.2 Write Critical Design Review David Nov. 2, 2016 Dec. 9, 2016 1.4 Flight Readiness Review (Report) David - - - 1.4.1 Create Standards for Flight Readiness Review David Jan. 1, 2017 Jan. 18, 2017 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 1.7.2 Preliminary Design Review Practice David Oct. 28, 2016 Oct. 28, 2016 1.8 Critical Design Review (Presentation) David - - - 1.8.1 Create Critical Design Review Presentation David Jan. 1, 2017 Jan. 11, 2017 1.8.2 Critical Design Review Practice David Jan. 11, 2017 Jan. 11, 2017 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. 27, 2016 Sept. 26, 2016 1.11.1 Budget Monitoring David - - - 1.12 Create Schedule David May. 25, 2016 Jun. 1, 2016 Aug. 25, 2017 1.13 Create Detailed Task Breakdown David May. 1, 2016 Jun. 1, 2016 May. 1, 2016 1.14 Integration of Subsections David - - - 1.15 Create and Maintain Website Bryan Sept. 12, 2016 Sept. 16, 2016 1.16 Travel Arrangements for Testing & Competition David 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 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
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2 Propulsion Andrew 2.1 Motor Type Selection (General, Proposal Level) Andrew Sep. 16, 2016 2.1.1 Motor Research Andrew 1-Jul Aug. 19, 2016 Aug. 19, 2016 2.1.2 Motor Comparision Andrew 1-Jul Sept. 14, 2016 Sept. 13, 2016 2.1.3 Motor Elimination Andrew 1-Jul Sept. 14, 2016 Sept. 13, 2016 2.1.4 Caclulate projected Altitude Andrew - - - 2.1.5 Select projected motor Andrew Sept. 14, 2016 Sept. 15, 2016 Sept. 13, 2016 2.2 Mission Performance Predictions Andrew - 2.2.2 Simulated Thrust Curve Andrew Sept. 14, 2016 Sept. 15, 2016 Sept. 13, 2016 2.3 Conceptual Model Creation Andrew - 2.3.1 Motor Mount Design Andrew Aug. 15, 2016 Sep. 4, 2016 Sept. 19, 2016 2.3.1.1 Motor Fastening Design Andrew Aug. 15, 2016 Aug. 28, 2016 Sept. 19, 2016 2.3.1.2 Motor Placement Andrew Aug. 15, 2016 Sep. 4, 2016 Sept. 4, 2016 2.3.1.3 Redesign Andrew Sept. 5, 2016 Nov. 30, 2016 2.3.2 Rear Aerodynamics Design Andrew 2.3.2.1 Collaboration with Aerodynamics Andrew Aug. 15, 2016 Nov. 30, 2016 2.3.4 Ignition Design Andrew 2.3.4.1 Ignition Research Andrew Sept. 15, 2016 Sept. 21, 2016 2.3.4.2 Ignition Placement Andrew Sept. 15, 2016 Sept. 21, 2016 2.3.4.3 Ignition Fastening Design Andrew Sept. 15, 2016 Sept. 21, 2016 2.3.4.4 Ignition Safety Interlock Design Andrew Sept. 15, 2016 Sept. 21, 2016 2.3.4.5 Igniter Installation Hatch Design Andrew Sept. 15, 2016 Sept. 21, 2016 2.3.4.6 Launch Switch w/ Returning to "off" Position Andrew Sept. 15, 2016 Sept. 21, 2016 2.3.4.4 Redesign Andrew Sept. 15, 2016 Nov. 30, 2016 2.4 Rocksim Modeling Andrew - 2.4.1 Model Rocket with Motor w/ Different Weights Andrew Aug. 15, 2016 Jan. 15, 2017 2.4.1.1 Simulation 1 Andrew Aug. 30, 2016 Sept. 14, 2016 Sept. 4, 2016 2.4.1.2 Discussion with Other Sections Andrew 15-Sep Sept. 21, 2016 Sept. 6, 2016 2.4.1.2 Resimulate Andrew Sept. 21, 2016 Sept. 30, 2016 Sept. 19, 2016 2.4.2 Simulate Full Scale Model Andrew 2.4.2.1 Preliminary Motor Selection Simulation Andrew Aug. 15, 2016 Sept. 14, 2016 Sept. 13, 2016 2.4.2.2 Preliminary Weighted Sections Simulation Andrew Aug. 15, 2016 Sept. 14, 2016 Sept. 13, 2016 2.4.2.3 Redesign Andrew Sept. 14, 2016 Sept. 21, 2016 Sept. 19, 2016 2.4.2.4 Final Motor Selection Simulation Andrew Sept. 15, 2016 Sept. 21, 2016 Sept. 13, 2016 2.4.2.5 Second Weighted Section Simulation Andrew Sept. 21, 2016 Sept. 25, 2016 Sept. 19, 2016 2.4.2.6 Redesign 2 Andrew Sept. 25, 2016 Sept. 29, 2016 Sept. 19, 2016 2.4.2.7 Final Rocket Simulation Andrew Sept. 29, 2016 Jan. 15, 2017 2.4.3 Simulate Half Scale Model Andrew 2.4.3.1 Physical Similitude Calculations Andrew Sept. 14, 2016 Nov. 30, 2016 2.5 Preliminary Design Review Andrew 2.5.1 Baseline Motor Selection Andrew Sept. 15, 2016 Sept. 16, 2016 Sept. 13, 2016 2.5.2 Thrust-Weight Ratio Andrew Sept. 15, 2016 Sept. 16, 2016 Sept. 13, 2016 2.5.3 Rail Exit Veloctiy Andrew Sept. 15, 2016 Sept. 16, 2016 Sept. 13, 2016 2.6 Critical Design Review David 2.6.1 Specify Motor Andrew Sept. 21, 2016 Oct. 7, 2016 2.6.2 Final Drawings Andrew Sept. 21, 2016 Oct. 7, 2016 2.6.3 Final Analysis and Model Results Andrew Sept. 29, 2016 Dec. 5, 2016 2.6.4 Motor Mounts Andrew Sept. 5, 2016 Nov. 30, 2016 2.6.5 Altitude Predictions with Final Design Andrew Sept. 29, 2016 Dec. 5, 2016 2.6.6 Actual Motor Thrust Curve Andrew Sept. 29, 2016 Dec. 5, 2016 2.6.7 Show Scale Model Results Andrew Sept. 29, 2016 Nov. 30, 2016 2.7 Critical Design Review Presentation David 2.7.1 Final Motor Choice Andrew Sept. 15, 2016 Oct. 7, 2016 2.7.2 Rocket Flight Stability in Static Diagram Andrew Sept. 15, 2016 Oct. 7, 2016 2.7.3 Thrust-to-Weight ratio Andrew Sept. 15, 2016 Oct. 7, 2016 2.7.4 Rail Exit Velocity Andrew Sept. 15, 2016 Oct. 7, 2016 2.8 Flight Readiness Review Presentation David 2.8.1 Final Motor Choice/ description Andrew Sept. 15, 2016 Oct. 7, 2016 2.8.2 Key Design Features Andrew Sept. 21, 2016 Nov. 30, 2016 2.8.3 Rocket Flight Stability Andrew Sept. 15, 2016 Oct. 7, 2016 2.8.4 Launch Thrust-Weight Ratio Andrew Sept. 15, 2016 Oct. 7, 2016 2.8.5 Rail Exit Velocity Andrew Sept. 15, 2016 Oct. 7, 2016 2.9 Testing Andrew 2.9.1 Ignition Testing Andrew Nov. 1, 2016 Feb. 12, 2017 2.9.1.1 Switch Testing Andrew Nov. 1, 2016 Feb. 12, 2017 2.9.1.2 Fuel Igition Testing Andrew Nov. 1, 2016 Feb. 12, 2017 2.9.1.3 Ignition Mount Testing Andrew Nov. 1, 2016 Feb. 12, 2017 2.9.1.4 Ignition Safety Interlock Testing Andrew Nov. 1, 2016 Feb. 12, 2017 2.9.1.5 Misfire Testing Andrew Nov. 1, 2016 Feb. 12, 2017 2.9.2 Motor Testing Junior Nov. 1, 2016 Feb. 12, 2017 2.9.2.1 Impulse Testing Junior Nov. 1, 2016 Feb. 12, 2017 2.9.2.1.1 Testing Junior Nov. 1, 2016 Feb. 12, 2017 2.9.2.1.2 Data Analysis Junior Nov. 1, 2016 Feb. 12, 2017 2.9.2.2 Thrust Testing Junior Nov. 1, 2016 Feb. 12, 2017 2.9.2.2.1 Testing Junior Nov. 1, 2016 Feb. 12, 2017 2.9.2.2.2 Data Analysis Junior Nov. 1, 2016 Feb. 12, 2017 2.9.2.4 Pressure Testing Junior Nov. 1, 2016 Feb. 12, 2017 2.9.2.2.1 Testing Junior Nov. 1, 2016 Feb. 12, 2017 2.9.2.2.2 Data Analysis Junior Nov. 1, 2016 Feb. 12, 2017 2.9.2.4 Motor Mount Testing Andrew Nov. 1, 2016 Feb. 12, 2017 2.9.3 FEA on Motor Mount Andrew 2.9.3.1 Vibration Analysis Andrew 2.9.3.2 Combustion Analysis Andrew 2.9.3.3 Modal Analysis Andrew 2.9.3.4 Stiffness Analysis Andrew 2.9.3.5 Impulse Analysis Andrew 2.9.3.6 Shear Stress Calculations Andrew 2.9.3.7 Shear Stress Analysis with FEA Andrew
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3 Aerodynamics Torsten 3.1 3D Modeling - Entire Rocket Torsten 1-May Oct. 26, 2016 3.1.1 General, Proposal-Level Rocket Model & Component Selection Torsten 1-May Sep. 30, 2016 3.1.2 Integration of Subcomponent Models into 3D Model Torsten 1-Aug Oct. 26, 2016 3.1.3 1/2 Scale 3D Model Torsten 1-Nov Nov. 20, 2016 3.1.4.Wind Tunnel Scale 3D Model Torsten 1-Feb Mar. 5, 2017 3.2 Fins, Body, Nose Cone Selection Torsten Oct. 9, 2016 3.2.1 Full Scale Selection Torsten 1-May Sep. 30, 2016 3.2.2 1/2 Scale Selection Torsten 1-Nov Nov. 20, 2016 3.2.3 Wind Tunnel Scale Selection Torsten 30-Mar Mar. 5, 2017 3.3 Fins, Body, Nose Cone Construction Torsten Jan. 22, 2017 3.2.1 Full Scale Construction Torsten 12-Feb Jan. 22, 2017 3.2.2 1/2 Scale Construction Torsten 30-Nov Dec. 4, 2016 3.2.3 Wind Tunnel Scale Construction Torsten 12-Jan Apr. 2, 2017 3.4 Paint Torsten 3.4.1 Paint Effect on Drag Torsten 1-Aug Oct. 26, 2016 3.4.2 Painting Torsten Not happening Jan. 22, 2017 3.5 Determination of Center of Mass Torsten 1-Aug Jan. 22, 2017 3.6 Determination of Center of Pressure Torsten 1-Aug Jan. 22, 2017 3.7 Optimization of Center of Mass vs Center of Pressure Torsten 1-Aug Jan. 22, 2017 3.8 CFX Modeling Torsten Jan. 15, 2016 3.8.1 Full Scale Rocket Performance Torsten 1-May Sep. 30, 2016 3.8.2 1/2 Scale Rocket Performance Torsten 1-Nov Nov. 20, 2016 3.8.3 Wind Tunnel Scale Performance Torsten 12-Jan Mar. 5, 2017 3.9 Collaboration with Launch Pad for Guides Torsten 1-Nov Jan. 22, 2017 3.10 Study Feasability of Real-Time Drag Changing Torsten 1-Aug Sep. 30, 2016 3.11 Redesign of Rocket Body, Nosecone, Fins Torsten 1-Nov Jan. 22, 2017
4 Payload A 4.1 Payload A Design Justin Aug. 20, 2016 Sept. 20, 2016 4.1.1 Official Altimeter Justin Aug. 20, 2016 Sept. 20, 2016 4.1.2 Radio Frequency and GPS Tracking Justin Aug. 20, 2016 Sept. 20, 2016 4.1.3 Arming and Disarming Electronics Justin Aug. 20, 2016 Sept. 20, 2016 4.2 Payload A Construction Justin Nov. 1, 2016 Nov. 20, 2016 4.2.1 Official Altimeter Justin Nov. 1, 2016 Nov. 20, 2016 4.2.2 Radio Frequency and GPS Tracking Justin Nov. 1, 2016 Nov. 20, 2016 4.2.3 Arming and Disarming Electronics Justin Nov. 1, 2016 Nov. 20, 2016 4.3 Payload A Redesign Justin Nov. 10, 2016 Nov. 20, 2016 4.4 Integration with Data Collection System Justin Aug. 20, 2016 Nov. 28, 2016 4.5 Data Transmission Justin - - 4.5.1 Wireless Receiver Justin Aug. 20, 2016 Nov. 1, 2016 4.5.1.1 Design Ground Station Wireless Receiver Justin Aug. 20, 2016 Nov. 1, 2016 4.5.1.2 Construct Ground Station Wireless Reciever Justin Nov. 1, 2016 Nov. 20, 2016 4.5.2 Wireless Transmission Justin Aug. 20, 2016 Nov. 20, 2016 4.5.1.1 Design Wireless Transmitter Justin Aug. 20, 2016 Nov. 20, 2016 4.5.1.2 Construct Wireless Transmitter Justin Nov. 1, 2016 Nov. 20, 2016 4.6 Create Test Plan & Test to Ensure Components in working order Justin Nov. 1, 2016 Dec.12, 2016 4.7 Collaboration with Payload B over Motherboard Justin - - 4.8 Determine if Separation is Necessary Justin Aug. 20, 2016 Sept. 20, 2016 4.9 Ensure that all components can be subjected to rocket stresses Justin Nov. 1, 2016 20-Jan 4.10 Meetings/Reports Justin - -
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5 Payload B Braden 5.1 Payload B Design (Fragile Material Housing) Braden 5.1.1 Design of Experiment Braden Aug. 1, 2016 Sep. 5, 2016 5.1.2 Design of Experimental Apparatus Braden Sep. 5, 2016 Ongoing and changing 5.1.3 Design of Mounting Braden Sep. 5, 2016 Sep. 20, 2016 5.2 Payload B Construction Braden 5.2.1 Construction of Experiment and housing Braden Oct. 1 Ongoing and changing 5.2.2 Construction of Mounting Braden Oct. 1 Nov. 20, 2016 5.3 Payload Testing and Experimentation Braden 5.3.1 Design Testing Plan Braden Sept. 10, 2016 Sep. 30, 2016 5.3.2 Carry Out Testing Braden Oct. 10, 2016 Dec. 4, 2016 5.3.3 Data Analysis Braden Dec. 4, 2016 Jan. 22, 2017 5.3 Payload B Redesign Braden Jan. 22, 2017 feb. 1, 2017 5.4 Create Test Plan to Ensure Hardware in Good Working Order Braden Sept. 10, 2016 Sep. 30, 2016 5.5 Collaboration with Payload A over Data Collection Braden Sep. 10, 2016 Sep. 30, 2016 5.6 Determine if Separation is Necessary Braden 10-Sep Sep. 30, 2016 5.7 Ensure that all components can be subjected to rocket stresses Braden Jan. 22, 2017 feb. 1, 2017 5.7 Reports Braden Sept. 15, 2016 5.8 Meetings/Group Work
6 Recovery Stewart 23-Jan 3-Feb 6.1 Recovery System Design Stewart 15-Aug 30-Sep 6.1.1 Recovery System Research Stewart 15-Aug 9-Sep 6.1.2 Recovery System Component Selection Stewart 29-Aug 30-Sep 6.1.2.1 Parachutes (Drogue & Main) Stewart 29-Aug 30-Sep 6.1.2.2 Altimeters Stewart 29-Aug 9-Sep 6.1.2.3 Shock cord and hardware Stewart 29-Aug 9-Sep 6.1.2.4 Ejection system Stewart 29-Aug 9-Sep 6.1.2.5 Bulkhead components Stewart 29-Aug 9-Sep 6.1.2.6 Electronics Stewart 29-Aug 9-Sep 6.1.3 Bulkhead design Stewart 29-Aug 30-Sep 6.1.4 Circuit design & programming Stewart 29-Aug 30-Sep 6.1.5 Computer Modeling 6.1.5.1 Parachute modeling Stewart 29-Aug 30-Sep 6.1.5.2 3D Assembly 6.1.5.3 Finite Element Analysis 6.1.6 Scaled model design Stewart 3-Oct 28-Oct 6.1.6.1 Parachutes (Drogue & Main) Stewart 29-Aug 30-Sep 6.1.6.2 Shock cord and hardware Stewart 29-Aug 9-Sep 6.1.6.3 Bulkhead components Stewart 29-Aug 9-Sep 6.1.6.4 Ejection system Stewart 29-Aug 9-Sep 6.2 Recovery System Construction Stewart 31-Oct 2-Dec 6.2.1 Bulkhead assembly Stewart 31-Oct 4-Nov 6.2.2 Circuit assembly Stewart 7-Nov 11-Nov 6.2.3 Ejection system assembly Stewart 14-Nov 18-Nov 6.2.4 Full-system integration Stewart 21-Nov 2-Dec 6.2.5 Scaled model construction Stewart 31-Oct 2-Dec 6.3 Recovery System Testing Stewart 5-Dec 3-Feb 6.3.1 Parachute testing (multiple wind speeds) Stewart 5-Dec 3-Feb 6.3.2 Ejection system testing Stewart 9-Jan 20-Jan 6.3.3 Circuit and transmitter testing Stewart 9-Jan 20-Jan 6.3.4 Full-system testing Stewart 23-Jan 3-Feb 6.4 Launch Pad David 6.4.1 Launch Pad Design David Sept. 30, 2016 6.4.2 Launch Pad Material Aquisition David Oct. 10, 2016 6.4.3 Launch Pad Fabrication David Oct. 25, 2016 6.5 Obtain Launch License Stewart 4-Nov 4-Dec
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7 Testing Bryan 7.1 Oversee all Subsection Testing Bryan Dec. 12, 2016 5-Apr 7.2 Manage Junior Level Testing Bryan Dec. 12, 2016 17-Mar 7.3 1/2 Scale Testing Bryan - - 7.3.1 Design of 1/2 Scale Testing Experiments Bryan Sept. 30, 2016 Dec. 2, 2016 7.3.2 Construction and Conduction of 1/2 Scale Testing Experiments Bryan Dec. 2, 2016 Dec. 7, 2016 7.3.3 Assess CFX with Results Bryan Jan. 9, 2017 Jan. 14, 2017 7.4 Wind Tunnel Testing Bryan Feb. 5, 2017 Feb. 26, 2017 7.4.1 Assess CFX with Results Bryan 20-Mar 25-Mar 7.5 Work with Subsections to Optomize Sections based on Testing Bryan Dec. 12, 2016 25-Mar 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 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. 9, 2016
8 Safety Bryan 8.1 Create a Detailed Step-by-Step Launch Procedure Bryan Nov. 7, 2016 Dec. 8, 2016 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 Dec. 2, 2016 8.2 Designated Head of Safety Bryan - - 8.3 Creation of Safety Checklist Bryan Aug. 29, 2016 Sept. 30, 2016 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 Oct. 28, 2016 9.2 Create Report for Educational Engagement Activity Bryan Nov. 7, 2016 Nov. 11, 2016 9.3 Create Presentation for Educational Engagement Activity Bryan Nov. 14, 2016 Nov. 18, 2016 9.4 Create Display for Educational Engagement Activity Bryan Nov. 14, 2016 Nov. 18, 2016 Nov. 28, 2016 Dec. 2, 2016
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