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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

34 | P a g e

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.

36 | P a g e

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

37 | P a g e

of thread for the cap to screw down at variable lengths. The reason for this is that Cylinder 1

will contain support material (material to be determined through testing) and when the unknown

object(s) are placed within the container, the support material will be displaced the same volume

as the object(s). To be sure the support material firmly holds the object(s), the lid will screw to

variable distances to compress the material and object(s) regardless of their size.

Figure 23 – Payload Inner Cylinder

Attached to Cylinder 2 are 5 base springs - designed to absorb most of the large impact

forces such as the initial takeoff, parachute deployment, and landing. Prior to completing the

first mathematical model utilizing Simulink, these springs were not included. However, the

forces induced on Cylinder 1 and thus the fragile material were too large so a series and parallel

spring system was created by introducing the 5 base springs. These springs can be seen in Figure

24 and are labeled a.

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Figure 24 - Springs Attached to the Main Payload and Bulkhead Attachment

Then entire system consisting of Cylinder 1, Cylinder 2, 12 wire rope isolators, and 5 base

springs, oscillates within the body tube of the rocket and is mounted to the bulkhead separating

the payload bay and the recovery bay. This bulkhead can be seen in Figure 24 is labeled “b”.

The walls of the payload bay, as well as the outside of Cylinder 2, will be lubricated to ensure

smooth translation during oscillation with graphite powder. Again, the exploded view of the

entire system can be seen in Figure 21.

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.

44 | P a g e

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

58 | P a g e

_________

_________

_________

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.

59 | P a g e

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

64 | P a g e

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.

76 | P a g e

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