university of evansville student launchi | p a g e enclosed: preliminary design review submitted by:...
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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
University of Evansville Student Launch
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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.
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.
Main Payload
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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.
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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
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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 X О О Carbon Fiber provides the highest
tensile strength and lowest
ductility at the highest cost
Fiberglass Δ Δ Δ Fiberglass provides a moderate
strength, ductility, and cost
relative to Carbon Fiber and Blue
Tube.
Blue Tube О X X Blue Tube provides the lowest
ductility and strength at the
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
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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 X О О Carbon Fiber provides a high
tensile strength and low ductility
at a high cost.
Fiberglass О Δ Δ Fiberglass provides a moderate
strength, ductility, and costs
significantly less than Carbon
Fiber or ULTEM.
ULTEM X О О ULTEM provides a high tensile
strength and low ductility at a
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 X О Δ Aluminum offers the highest
strength of all materials
considered. It comes at an
increased cost and weight.
Plywood O X O Plywood offers the lowest cost
and weight at the price of
strength.
Fiberglass Δ Δ Δ Fiberglass offers a moderate
alternative to plywood and
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
Option Stability Ease of
Manufacturing
Likelihood
of Damage Overview Decision Explanation
Clipped Delta Δ О O The Clipped Delta is the easiest to
manufacture and offers moderate
stability and drag.
Trapezoidal X Δ О The Trapezoidal offers the lowest
drag but the least stability.
Tapered Swept О Δ X The Tapered Swept offers the
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
Ogive Δ О The Ogive nosecone is the most difficult to
manufacture and thus the most expensive but
offers the lowest drag.
Elliptical O Δ The Elliptical nosecone can be purchased at a
moderate cost for a moderate drag.
Conical O X The Conical nosecone is the easiest to
manufacture and thus the least expensive but
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
Option Cost Safety Against
Regulations Overview Decision Explanation
Cluster Motor
Mount O Δ X The cluster motor would cost less
and reach the optimal altitude
with minimal safety concerns.
Single Motor
Mount X Δ O The single motor cost is high and
creates concerns about safety and
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 O X О Plywood is great for weight and
cost but the strength is a problem
for large motors
G10 Fiberglass Δ О О The cost of fiberglass is budget-
able because of the high strength
and the weight of the material
Aluminum Δ О X Aluminum has a good cost
associated with machining it in
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
Stratologger
CF O O Δ
For a low cost, this altimeter
provides a full set of features with
a higher power draw.
AltusMetrum
EasyMini Δ Δ О For a medium cost, this altimeter
provides a reduced feature set
with a low power draw.
Entacore
AIM3 X O О For a high cost, this altimeter
provides a full set of features with
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
Elastic O X X For a very low cost, elastic
provides a low-strength, high-
elasticity solution.
Tubular Kevlar X O Δ For a high cost, Kevlar provides
the greatest strength with a very
low elasticity.
Tubular
Nylon Δ Δ О For a medium cost, nylon
provides acceptable strength at a
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
21 | 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
Option Cost Descent
Rate Max Force Overview Decision Explanation
24” Fruity
Chutes Classic
Elliptical O Δ X
For a low cost, a relatively quick
descent rate can be achieved, but
at the cost of a large maximum
force at main parachute
deployment.
36” Fruity
Chutes
Classic
Elliptical
O O Δ
For a low cost, a good descent
rate can be achieved with an
acceptable maximum force at
main parachute deployment.
48” Fruity
Chutes Classic
Elliptical Δ Δ О
For a medium cost, a relatively
slow descent rate can be achieved
with a low maximum force at
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
Option Cost
Ground
Impact
Velocity
Max Force Overview Decision Explanation
72” Fruity
Chutes Iris
Ultra O X O
For a low cost, an unacceptable
ground impact velocity can be
achieved with a good maximum
force.
84” Fruity
Chutes Iris
Ultra O Δ Δ
For a low cost, an acceptable
ground impact velocity can be
achieved with an acceptable
maximum force.
96” Fruity
Chutes Iris
Ultra Δ O Δ
For a medium cost, a good impact
velocity can be achieved with an
acceptable maximum force.
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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Figure 8 - Propulsion Components Labeled
Figure 9 - Dimensional Drawing for the Motor Mount
Bulkhead Centering Rings Thrust Plate
Inner Tube Motor and Case
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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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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.
0.84”
2”
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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
Figure 14 - Recovery system layout within airframe
12” 9” 14”
35”
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Figure 15 – Tethering of Rocket Sections
25’ 35’
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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 17 - Exploded View of Electronic Payload
Figure 16 - Electronic Payload within Nosecone
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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
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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.
39 | P a g e
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 1)
Weight per cubic ft. Density Grain Size Liquid/Solid Viscosity
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 system)
Varies 3 lb/ft3 N/A Solid N/A
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.
40 | 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
42 | P a g e
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
43 | 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
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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
46 | 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
47 | 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.
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.
0
1000
2000
3000
4000
5000
6000
0 20 40 60 80 100 120 140 160 180
Alt
itu
de
(ft)
Time (s)
48 | P a g e
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.
49 | P a g e
% 𝐷𝑖𝑓𝑓𝑒𝑟𝑒𝑛𝑐𝑒 = 𝑂𝑝𝑒𝑛𝑅𝑜𝑐𝑘𝑒𝑡 𝑉𝑎𝑙𝑢𝑒−𝑅𝑜𝑐𝑘𝑠𝑖𝑚 𝑉𝑎𝑙𝑢𝑒
𝑂𝑝𝑒𝑛𝑅𝑜𝑐𝑘𝑒𝑡 𝑉𝑎𝑙𝑢𝑒 Eq. 1
% 𝐷𝑖𝑓𝑓𝑒𝑟𝑒𝑛𝑐𝑒 = 5379𝑓𝑡−5368𝑓𝑡
5379𝑓𝑡𝑥 100 = 0.20% Eq. 2
Figure 32 - Predicted Altitude from Rocksim Simulation
0
1000
2000
3000
4000
5000
6000
0 20 40 60 80 100 120 140 160 180
Alt
itu
de
(ft)
Time (s)
50 | P a g e
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
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Figure 35 - Thrust Curve for the L850W Motor in OpenRocket
Figure 36 - Thrust Curve for the L850W Motor in Rocksim
0
200
400
600
800
1000
1200
1400
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
Thru
st (
N)
Time (s)
-200
0
200
400
600
800
1000
1200
1400
0 1 2 3 4 5 6
Thru
st (
N)
Time (s)
52 | P a g e
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.
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 &
Payload
9.19 20.9
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
CP CG
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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
__________________________________
57 | P a g e
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
Epoxy
Inhalation of toxic fumes, accidental
ingestion, or contact with skin leading
to irritation or rash
Minor High Work in well ventilated
spaces
Dust
Particles
Inhalation of dust particles from
sanding or machining operations
resulting in breathing problems
Minor High
Wear mask when sanding
to avoid inhaling dust
particles
Heavy Tools
and
Machinery
in Lab
Improper handling of shop tools or
machining operations leading to
personal injury or destruction of
equipment
Significant Medium
Ensure proper training for
all team members working
with any tool or machinery
in shop
Rocket
Propellant
Exposure to rocket fuel in contact with
skin leading to irritation and burns Major Medium
Properly transport motor
from offsite location to
launch site
Black
Powder
Gases may be toxic if exposed in areas
with inadequate ventilation. Also keep
away from open flame, sparks, and heat
Major Low
Store in portable fireproof
case to keep away from
fire and high temperatures
Craft and
Exacto
Knives
Cuts leading to injury as a result of
precision cutting operations on fins or
other pieces of the rocket body
Minor Medium
Ensure at least one
teammate is working
alongside the person doing
the cutting. Practice safe
cutting procedures by
cutting away from body.
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Fire
Burns, significant and/or fatal injury, or
damage to school from fire as a result
of faulty wiring, or improper handling
of the motor and black powder
Major Low
Store a fire extinguisher in
the room where the rocket
will be constructed. If an
object starts to overheat, let
it cool and have the fire
extinguisher ready
Handheld
Tools
Bruises, cuts or scrapes from
mishandling of basic handheld shop
tools such as hammer or saw
Significant High
Be aware of surroundings
when operating the
handheld tools and ensure
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
Risk/Hazard Effect Severity Likelihood Proposed
Mitigation
Motor
Handling/Accidental
Ignition
Improper handling or storage of
motor resulting in accidental or
unexpected ignition
Major Low
Properly transport
motor from offsite
location to launch
site. Ensure proper
connections before
launch
Launch Failure Failure of motor to ignite and
launch rocket properly Significant Low
Maintain safe
distance from
launch pad. Have
team mentor/safety
officer inspect
rocket on launch
pad
Main Parachute
Deployment Failure
Failure of the secondary parachute
to deploy leading to freefall or
unstable flight of rocket back to
the ground
Major Low
Maintain safe
distance from
launch pad
Drogue Parachute
Deployment Failure
Failure of the initial parachute to
deploy leading to freefall or
unstable flight of rocket back to
the ground
Significant Low
Maintain safe
distance from
launch pad
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Instability During
Flight
Failure of the rocket to maintain
its projected flight path due to
unforeseen design flaw or in flight
malfunction
Major Low
Maintain safe
distance from
launch pad
Altimeter or Other
Electronics in
Avionics Bay
Malfunction/Fall
Off
Potential short circuiting or harm
to spectators below Minor Medium
Verify all
electronics work
properly before
launch and are
firmly attached to
the rocket
Coupler Excessively
Tight
Failure of parachute to deploy
leading to damage to rocket Major Low
Run multiple tests
to ensure proper
amounts of black
powder is used to
allow rocket to
separate
Payload Not
Secured Properly
Inability to return materials
without breaking Minor Medium
Take caution when
inputting payload
into rocket before
launch and ensure
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
Epoxy Fumes
When epoxying various
pieces of the rocket
together, harmful fumes
are released into the
atmosphere
Minor High
Work in well ventilated
spaces and dispose of
waste properly
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Dust Particles
Small dust particles from
sanding or machining
operations are released into
the environment which can
result in breathing
problems
Minor High
Wear mask when sanding
to avoid inhaling dust
particles and try to
contain dust when
sanding opposed to freely
releasing it into
surrounding air.
Rocket Motor
Ignition
Upon ignition, the motor
reaches high temperatures
and hot exhaust is released,
which could potentially
burn the areas where the
rocket is launched or lands
Major Low
Place flame resistant
material beneath the
launch pad to avoid
burning the immediate
surroundings or starting a
fire
Debris from Rocket
If various pieces of the
rocket do not stay intact
during decent, or the
parachutes do not operate
properly, pieces of the
rocket could break off
during flight or upon
impact and be irretrievable,
leading to minor
environmental harm.
Significant Low
Ensure fully functioning
parachutes before launch
via pre-launch checklist
and check that all
components of the rocket
and payload are
accounted for upon
return.
Environmental Effects on Vehicle
Water
Precipitation and moisture
within the rocket could
affect the structural
integrity of the rocket, or
could lead to malfunctions
of the electronics housed in
the avionics bay
Significant Low
Avoid launching rocket
in wet conditions and
ensure a dry area for
storage and transport
Wind
Strong wind or
unpredictable wind gusts
can cause the rocket to
deviate from its ideal flight
path and can lead to
damage to the rocket and
potential harm to spectators
Significant Medium
Avoid launching rocket
on days where high speed
winds or unpredictable,
strong wind gusts are
present
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Humidity
Humidity can lead to
moisture in the body of the
rocket which can lead to
corrosion and weakening
of various materials used to
construct the rocket. It can
also negatively impact on-
board electronics
Minor Low
Store rocket in a dry area
to avoid moisture
entering the rocket over
time via humid air
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
Value
Likelihood Proposed Mitigation
Limited
Resources
Due to the new nature of the
project to this team specifically, if
the team is unable to find
valuable insight from external
sources, the design and
performance of the rocket could
suffer
High Medium
The team will work with
faculty members as well
as local rocketry club
members in order to gain
a better understanding of
rocketry and develop a
functional rocket.
Tight or
Minimal Budget
Lack of flexibility in the budget
could lead to the team being
forced to use parts that are not
optimal, or being unable to
replace parts of the rocket that are
broken during testing
High High
The team and its adult
educators will apply for
grants and fundraise to
provide the team with a
flexible budget beyond
the normal amount of
money allotted to the
project by the school
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Mismanagement
of Time
Inability of the team to keep up
with the initial schedule set forth
in the task breakdown could lead
to major delays, poor quality of
work, or the rocket not being
completed by competition
Medium Low
Team members will fill
out weekly time cards
and log their hours in the
task breakdown in order
to ensure everyone
remains on schedule
Underestimation
of Scope of
Work
Failure to properly account for
the work needed to complete the
project could lead to the project
running behind schedule and
various facets of the rocket not
being completed in a quality
manner
Medium Low
There will be constant
communication amongst
all team members and
with NASA to ensure the
scope of work is clear
Increase in
Safety
Regulations
Adding material to the rocket in
order to increase safety will result
in an increase in expenses
Low Medium
The team will design and
downselect with safety in
mind, and will clearly
identify all safety
measures before
construction so that
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
Number
Summarized
Requirement
Verification
Method(s) Description of Verification Plan
1.1
The vehicle shall
deliver the science or
engineering payload
to an apogee altitude
of 5,280 feet above
ground level (AGL).
Test
Analysis
The rocket team will utilize OpenRocket,
RockSim, CFD, & test flight data to
achieve an accurate prediction of altitude.
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NASA Requirements
Handbook
Number
Summarized
Requirement
Verification
Method(s) Description of Verification Plan
1.2
The vehicle shall carry
one commercially
available, barometric
altimeter for recording
the official altitude
used in determining
the altitude award
winner.
Inspection
The rocket will house a Atlus Metrum
TeleMega altimeter in the nosecone to
record the official altitude used in
determining the altitude award winner.
1.3
All recovery
electronics shall be
powered by
commercially
available batteries.
Inspection Batteries & altimeter will be purchased
from online rocketry sources.
1.4
The launch vehicle
shall be designed to be
recoverable and
reusable. Reusable is
defined as being able
to launch again on the
same day without
repairs or
modifications.
Test
Inspection
The rocket is reusable in design because
our team is using a motor that has refuels
that can be reloaded into the motor under
supervision.
1.5
The launch vehicle
shall have a maximum
of four (4)
independent sections.
Inspection
The launch vehicle will have 3
independent sections: the aft body tube,
the bow body tube and nosecone, and the
coupler.
1.6
The launch vehicle
shall be limited to a
single stage.
Inspection
Demonstration The launch vehicle shall be a single stage.
1.7
The launch vehicle
shall be capable of
being prepared for
flight at the launch
site within 4 hours.
Inspection
Demonstration
The launch vehicle will be designed with
an efficient and quick to construct design
that requires fewer than 4 hours to
prepare.
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NASA Requirements
Handbook
Number
Summarized
Requirement
Verification
Method(s) Description of Verification Plan
1.8
The launch vehicle
shall be capable of
remaining in launch-
ready configuration at
the pad for a
minimum of 1 hour
without losing the
functionality of any
critical on-board
component.
Test The launch vehicle design will ensure all
components have a life of greater than 1
hour without loss of functionality.
1.9
The launch vehicle
shall be capable of
being launched by a
standard 12-volt direct
current firing system.
Inspection
Test
The ignition system will be using a 12
volt direct current firing system.
1.10
The launch vehicle
shall require no
external circuitry or
special ground support
equipment to initiate
launch (other than
what is provided by
Range Services).
Inspection
There will be no external circuity for the
ignition system because it will be a
ground based ignition system being
placed underneath the rocket before
launch with 300 ft of cord between the
igniter and the controller.
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NASA Requirements
Handbook
Number
Summarized
Requirement
Verification
Method(s) Description of Verification Plan
1.11
The launch vehicle
shall use a
commercially
available solid motor
propulsion system
using ammonium
perchlorate composite
propellant (APCP)
which is approved and
certified by the
National Association
of Rocketry (NAR),
Tripoli Rocketry
Association (TRA),
and/or the Canadian
Association of
Rocketry (CAR).
Inspection
The motor being used is a solid fuel
motor from AeroTech. The motor is the
L850W.
1.12
Pressure vessels on
the vehicle shall be
approved by the RSO.
Inspection No pressure vessels will be used.
1.13
The total impulse
provided by a
University launch
vehicle shall not
exceed 5,120 Newton-
seconds (L-class).
Test
Analysis
The motor will produce an impulse of
3695 N-s which is below the specified
total impulse that is allowed.
1.14
The launch vehicle
shall have a minimum
static stability margin
of 2.0 at the point of
rail exit.
Test
Analysis
The launch vehicle will have a static
stability margin of 2.67.
1.15
The launch vehicle
shall accelerate to a
minimum velocity of
52 fps at rail exit.
Test
Analysis
The rocket team will utilize OpenRocket,
RockSim, CFD, & test flight data to
achieve an accurate prediction of
minimum velocity at rail exit. The current
value is 67.2 ft/s.
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NASA Requirements
Handbook
Number
Summarized
Requirement
Verification
Method(s) Description of Verification Plan
1.16
All teams shall
successfully launch
and recover a subscale
model of their rocket
prior to CDR.
Test
Analysis
A subscale model with comparable
weights, lengths, and masses will be
launched prior to the CDR.
1.17
All teams shall
successfully launch
and recover their full-
scale rocket prior to
FRR in its final flight
con- figuration.
Test
Analysis
The project schedule will ensure a full-
scale rocket launch occurs before the
FRR.
1.18
Any structural
protuberance on the
rocket shall be located
aft of the burnout
center of gravity.
Test
Analysis
The rocket will have 3 bolts holding the
nosecone to the bow body tube and shear
pins holding the coupler to the bow and
aft body tubes. These structural
protuberances are all located aft of the
burnout center of gravity
1.19 Vehicle Prohibitions
Inspection
Test
Analysis
The launch vehicle will follow all
prohibitions laid out in section 1.19 of the
2017 SL NASA Student Handbook.
2.1
Vehicle must deploy a
drogue parachute at
apogee, followed by a
main parachute at a
much lower altitude.
Demonstration
Inspection
Dual-deployment altimeters will be
programmed to fire ejection charges at
apogee and at ~1000 feet.
2.2
A successful ground
ejection test for both
parachutes must be
conducted prior to
sub- and full-scale
launches.
Test Multiple ejection tests will be conducted
prior to sub- and full-scale launches.
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NASA Requirements
Handbook
Number
Summarized
Requirement
Verification
Method(s) Description of Verification Plan
2.3
No part of the launch
vehicle may have a
kinetic energy of
greater than 75 ft-lbf
at landing.
Analysis
Demonstration
Parachute sizes will be optimized to
minimize kinetic energy at ground
impact.
2.4
Recovery electrical
circuits must be
independent of
payload circuits.
Inspection Recovery electronics will be located in a
separate, shielded coupler.
2.5
Recovery system must
include redundant,
commercial
altimeters.
Inspection Two PerfectFlite Stratologger CF
altimeters will be used.
2.6
Motor ejection cannot
be used for primary or
secondary
deployment.
Demonstration
Inspection
Black powder ejection charges will be
used to eject parachutes.
2.7
Each altimeter must
be armed by a
dedicated switch
accessible from the
rocket exterior.
Inspection Locking rotary switches and LED
indicators will be used to confirm the
state of the recovery electronics.
2.8
Each altimeter must
have a dedicated
power supply.
Inspection Separate 9-Volt batteries will be used to
power the altimeters.
2.9
Each arming switch
must be lockable to
the “ON” position.
Inspection Locking rotary switches will be used to
arm each altimeter.
2.10
Removable shear pins
must be used to seal
the parachute
compartments.
Inspection Threaded nylon shear pins will be used to
seal the parachute compartments.
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NASA Requirements
Handbook
Number
Summarized
Requirement
Verification
Method(s) Description of Verification Plan
2.11
Tracking device(s)
must transmit the
position of any parts
of the launch vehicle
to a ground receiver.
Test
Demonstration
Inspection
All parts of the launch vehicle will be
tethered together; position will be
transmitted via a flight computer in the
nosecone.
2.12
Recovery system
electronics must not
be adversely affected
by any other on-board
electronics.
Test
Inspection
Recovery electronics will be located in a
separate, shielded coupler.
3.4.1
Design container
capable of protecting
an unknown object of
unknown size and
shape.
Testing
Math model is used to develop spring
system in conjunction with a concentric
cylinder model to provide sufficient
vibration dampening and force reduction.
3.4.1.2 Object must survive
duration of flight Testing
The spring and concentric cylinder design
will be tested with a matrix of different
support materials as well as testing
materials to assure the unknown object(s)
can survive the flight during
demonstration.
3.4.1.4
Once the object is
obtained, it must be
sealed in its housing
until after the launch
and no excess material
may be added after
receiving the object.
Demonstration
Support material within cylinder 1 that
allows object to be inserted and not spill
any material such as a high viscosity fluid
or malleable solid.
4.1
Each team shall use a
launch and safety
checklist
Inspection
Final assembly and pre-launch checklists
will be created and reviewed at the
appropriate time to ensure safe launch of
the rocket and all members involved in
the launch
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NASA Requirements
Handbook
Number
Summarized
Requirement
Verification
Method(s) Description of Verification Plan
4.2
Each team shall
identify a student
safety officer who
shall be responsible
for the safety of the
team and ensure all
proper rules and
guidelines are
followed
Inspection
The team has appointed a safety officer to
monitor the safety of the team throughout
the project and ensure all federal rules
and laws are met.
4.3
The team safety
officer shall monitor
team activities with an
emphasis on safety
throughout the design,
construction, and
testing of the rocket
by maintaining MSDS
sheets and hazard
analyses
Inspection
The team safety officer will monitor the
progress of the project emphasizing the
proper safety procedures for the current
stage of the project.
4.4
Each team shall
appoint a mentor who
has certification and is
in good standing with
the NRA. This
member will be
designated as the
individual owner of
the rocket and
assumes liability
Inspection
The team has assigned an school faculty
member to mentor the project to provide
valuable insight on the rocket design and
construction as well as assume full
liability of the rocket.
4.5
During test flights,
teams shall abide by
the rules and guidance
of the local rocketry
club's RSO
Demonstration
Team will converse with RSO at local
rocketry club to ensure all of their
chapter’s rules and regulations are abided
by.
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NASA Requirements
Handbook
Number
Summarized
Requirement
Verification
Method(s) Description of Verification Plan
4.6
Teams shall abide by
all rules set forth by
the FAA
Demonstration
Team will converse with NASA lead
safety officer and thoroughly research all
rules and regulations set forth by the FAA
to ensure all rules and regulations are
abided by.
5.1
Students shall do
100% of the project
excluding motor /
black powder
handling.
Demonstration
Inspection
The team will continuously demonstrate
an independently managed and executed
project. The team lead will routinely
monitor this quality.
5.2 A detailed project plan
shall be maintained. Demonstration
Documents for scheduling, budget
tracking, outreach, and safety will be
continuously updated and reported.
5.3
Foreign National
members shall be
identified by the PDR.
Inspection The team lead will ensure that any
Foreign National members are clearly
indicated in the PDR.
5.4
All team members
attending launch week
shall be identified by
the CDR.
Inspection
It will be checked that a list of team
members, with indications of those
attending launch week, will be included
in the CDR.
5.5
The educational
engagement
requirement shall be
met by the FRR.
Inspection
The Educational Engagement lead shall
confirm that all documentation has been
received and approved by NASA prior to
the FRR.
5.6
The team shall
develop and host a
website for
documentation.
Test Team members will periodically confirm
that the website is functioning as intended
by opening each posted document.
5.7
The team shall post &
make available for
download all
deliverables by the
specified date.
Inspection The team lead shall confirm that all
documents are posted prior to the
specified date.
5.8 All deliverables must
be in PDF format. Inspection
The team lead shall confirm that all
documents posted are in PDF format.
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NASA Requirements
Handbook
Number
Summarized
Requirement
Verification
Method(s) Description of Verification Plan
5.9
A table of contents
must be included in all
reports.
Inspection The team lead shall ensure that a table of
contents is located at the start of each
report.
5.10
Page numbers shall be
provided in each
report.
Test Page numbers shall be checked to the
table of contents to ensure continuity
throughout the report.
5.11
The team shall
provide
videoconference
equipment needed for
reviews.
Demonstration
Test
Videoconference rooms will be reserved
and trialed immediately prior to each
design review.
5.12
All teams shall use
launch pads provided
by the SLS provider.
Demonstration The team shall design the rocket to utilize
1515 12’ launch rail.
5.13
The team must
implement the EIT
accessibility
standards.
Demonstration
If software or applications are created
(not planned) the team will abide by 36
CFR Part 1194. Otherwise, all
components containing software will be
checked to ensure compliance.
Team requirements have been developed in addition to the NASA requirements. These
can be seen in Table 24.
Table 24 - Team Requirements
Team Requirements
Number Requirement Verification
Method Description of Verification Method
1
All reports shall be
compiled at least three
days prior to NASA
due dates.
Demonstration
Reports shall be completed, according to
team schedule, prior to NASA due dates
to allow for revision time and mitigate
risk of late submissions.
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Team Requirements
Number Requirement Verification
Method Description of Verification Method
2
Each member of the
team shall have a
working knowledge of
each subsystem.
Inspection
At each team meeting, every sub-section
lead will review the status of their section
with the entire team. The team leader
will confirm that the information
presented is sufficient.
3
Safety shall be made
the team’s first
priority.
Test The safety officer will periodically ask
team members what the most important
aspect of the project is.
4 Altimeters shall be in
good working order. Test
All altimeters shall be flown on sub-scale
and full scale flight tests. Altitude
readings will be compared to confirm
consistency.
5
The tracking system
shall be in good
working order.
Test
The tracking system shall be flown on
the sub-scale and full scale flight tests.
This will be used to find the rockets thus
confirming its operation.
6
A solid output signal
must be given from
triggered altimeters.
Test
Analysis
All altimeters will be triggered while
voltage is read on the output. This output
will be read to confirm it is acceptable.
7 All circuits shall be
checked prior to use. Demonstration
All circuits will be confirmed at each
node to ensure connections.
8
Impulse for the
parachute deployment
shall be determined
experimentally.
Test
Analysis
The main parachute shall have an
apparatus (strain gauge) attached to it
that enables a force to be read as it opens
at high speed. This will cut down in the
large ambiguity that exists in estimating
an impulse value.
9
A spring constant for
parachute cords shall
be determined
experimentally.
Test
Analysis
The spring constant shall be determined
using forces related to what is
experienced with parachute opening.
This helps when estimating energy
absorption by the cord when the chute
opens.
10
Payload must reduce
force felt by object(s)
by 50 %
Testing From the mathematical model,
appropriate springs will be selected to
induce oscillation and reduce force.
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Team Requirements
Number Requirement Verification
Method Description of Verification Method
11
Payload must reduce
acceleration of
object(s) by 35 %
Testing
From the mathematical model,
appropriate springs will be selected to
insure acceleration graphs show 35
percent reduction from inputs.
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
Engagement
$100.00 1%
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
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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
Activity Due Date
NASA U.E. Team
Project Kickoff Aug. 15, 2016 - -
General Motor Selection/Data Sept. 30, 2016 - Sept. 16, 2016
Informal Design Sketches - Sept. 21, 2016 Sept. 19, 2016
Proposal Sept. 30, 2016 Oct. 3, 2016 Sept. 27, 2016
Motor Selection/ Data Oct. 31, 2016 Oct. 7, 2016
Proposal Presentation - Oct. 24, 2016 Oct. 22, 2016
PDR Report Nov. 04, 2016 - Oct. 26, 2016
PDR Flysheet Nov. 04, 2016 - Oct. 26, 2016
PDR Presentation Nov. 04, 2016 - Oct. 28, 2016
Sub-Scale Launch Motor Selection - - Nov. 30, 2016
Sub-Scale Launch - - Dec. 11, 2016
Design Report - Dec. 2, 2016 Nov. 29, 2016
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$
-$
-$
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$
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$
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 $8.98 8.98$
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$
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$
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$
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$
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$
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 End Date
Person
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
Start Date
Estimated ActualActualEstimated
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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