Critical Design Review Paper for Project A.R.E.S.

Georgia Institute of Technology Team ARES
120 North Ave NW Atlanta GA 30332
Project Hermes
January 15, 2016

2015-2016 Georgia Tech Team ARES
Critical Design Review
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Table of Contents
1. INTRODUCTION...............................................................................................................................4
1.1. TEAM SUMMARY ..........................................................................................................................4
1.2. WORK BREAKDOWN STRUCTURE .................................................................................................4
1.3. LAUNCH VEHICLE SUMMARY ......................................................................................................5
1.4. AGSE SUMMARY .........................................................................................................................5
2. CHANGES MADE SINCE PDR .......................................................................................................7
2.1. LAUNCH VEHICLE CHANGES........................................................................................................7
2.2. AGSE CHANGES...........................................................................................................................7
2.3. PROJECT PLAN CHANGES .............................................................................................................7
3. LAUNCH VEHICLE CRITERIA.....................................................................................................8
3.1. DESIGN AND VERIFICATION OF LAUNCH VEHICLE ......................................................................8
3.1.1. Mission Statement................................................................................................................8
3.1.2. Mission Success Criteria .....................................................................................................8
3.1.3. Major Milestone Schedule ...................................................................................................9
3.2. SYSTEM DESIGN REVIEW ...........................................................................................................10
3.2.1. Final Motor Selection........................................................................................................13
3.3. SYSTEM LEVEL FUNCTIONAL REQUIREMENTS ..........................................................................14
3.4. MANUFACTURING APPROACH....................................................................................................15
3.4.1. Payload Section .................................................................................................................15
3.4.2. Avionics Section.................................................................................................................18
3.4.3. Booster Section ..................................................................................................................20
3.4.4. Status and Plans of Remaining Manufacturing and Assembly..........................................22
3.4.5. Integrity of Design .............................................................................................................22
3.5. SUBSCALE FLIGHT RESULTS.......................................................................................................26
3.5.1. Flight Data.........................................................................................................................26
3.5.2. Result Discussion...............................................................................................................27
3.5.3. Impact on Full Scale..........................................................................................................28
3.6. RECOVERY SUBSYSTEM .............................................................................................................30
3.6.1. Recovery System Hardware...............................................................................................30
3.6.2. Electrical Hardware ..........................................................................................................30
3.6.3. Kinetic Energy Calculations..............................................................................................31
3.7. MISSION PERFORMANCE PREDICTIONS ......................................................................................32

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3.7.1. Validity of Analysis............................................................................................................32
3.7.2. Drag Assessment................................................................................................................33
3.8. FUTURE TESTING ........................................................................................................................34
3.9. LAUNCH PROCEDURES ...............................................................................................................35
3.9.1. Launch Checklist ...............................................................................................................35
3.10. SAFETY AND ENVIRONMENT..................................................................................................39
3.10.1. Overview............................................................................................................................39
3.10.2. Failure Modes and Analysis ..............................................................................................40
3.10.3. Personnel Hazards.............................................................................................................44
3.10.4. Environmental Concerns ...................................................................................................44
4. FLIGHT SYSTEMS..........................................................................................................................46
4.1. OVERVIEW ..................................................................................................................................46
4.2. SIMULINK DESIGN OVERVIEW....................................................................................................48
4.3. CONTROLLER DESIGN OVERVIEW..............................................................................................50
4.3.1. Control Technique .............................................................................................................50
4.3.2. Sliding Mode Control ........................................................................................................51
5. AGSE CRITERIA.............................................................................................................................53
5.1. AGSE SUMMARY .......................................................................................................................53
5.2. DESIGN REVIEW OF AGSE EQUIPMENT .....................................................................................54
5.2.1. Robotic Payload Delivery System (RPDS) ........................................................................54
5.2.2. Rocket Erector System (RES).............................................................................................57
5.2.3. MIS.....................................................................................................................................59
5.2.4. Electronics .........................................................................................................................62
5.3. MISSION SUCCESS CRITERIA AND FUNCTIONAL REQUIREMENTS..............................................64
5.4. MANUFACTURING APPROACH....................................................................................................65
5.4.1. Overall Manufacturing Approach .....................................................................................65
5.4.2. RPDS..................................................................................................................................65
5.4.3. RES ....................................................................................................................................66
5.4.4. MIS.....................................................................................................................................66
5.4.5. Electronics .........................................................................................................................67
5.5. REMAINING MANUFACTURING AND ASSEMBLY........................................................................67
5.6. AGSE ELECTRONICS ..................................................................................................................68
5.6.1. AGSE Electronics Overview..............................................................................................68

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5.6.2. Drawings and Schematics..................................................................................................69
5.6.3. Block Diagrams .................................................................................................................70
5.6.4. Batteries/Power .................................................................................................................71
5.6.5. Switch and indicator Wattage and Location .....................................................................71
5.7. COMPONENT TESTING ................................................................................................................72
5.7.1. Electronics Pause Button Test...........................................................................................72
5.7.2. RES Lifting Test .................................................................................................................73
5.7.3. RPDS Payload Insertion Test ............................................................................................73
5.8. INTEGRATION PLAN....................................................................................................................74
5.9. FAILURE AND SAFETY ANALYSIS...............................................................................................75
6. PROJECT PLAN ..............................................................................................................................77
6.1. BUDGET PLAN.............................................................................................................................77
6.2. FUNDING PLAN ...........................................................................................................................78
7. PROJECT SCHEDULE ...................................................................................................................80
7.1.1. Critical Path Chart: CDR to PLAR...................................................................................80
7.2. SCHEDULE RISK..........................................................................................................................83
8. EDUCATIONAL OUTREACH PLAN AND STATUS.................................................................85
8.1. OVERVIEW ..................................................................................................................................85
8.2. ATLANTA MAKER’S FAIRE.........................................................................................................85
8.3. FIRST LEGO LEAGUE ..................................................................................................................85
8.4. CEISMC GT...............................................................................................................................86
9. CONCLUSION..................................................................................................................................86
APPENDIX I: SCHEDULER ...................................................................................................................87
APPENDIX II: GANTT CHART.............................................................................................................90
APPENDIX III: TEST PLANS ................................................................................................................91
APPENDIX IV: CONTROL SYSTEM PAPER .....................................................................................92

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1. Introduction
1.1. Team Summary
Table 1: Team Summary
Team Summary
School Name Georgia Institute of Technology
Mailing Address North Avenue NW, Atlanta GA 30332
Team Name Team Autonomous Rocket Equipment System (A.R.E.S.)
Project Title Hermes
Launch Vehicle Name Skyron
Project Lead Victor R.
Safety Officer Stephen K
Team Advisors Dr. Eric Feron
NAR Section Primary: Southern Area Launch vehiclery (SoAR) #571
NAR Contact, Number &
Certification Level
Primary Contact: Joseph Mattingly
NAR/TRA Number: 92646
Certification Level: Level 2
Secondary: Jorge Blanco
1.2. Work breakdown structure
Team Autonomous Rocket Erector System (ARES) is composed of twenty-one students studying
varying fields of engineering. Our team is composed of less than 50% Foreign Nationals (FN) per
NASA competition requirements. To work more effectively, the team is broken down into groups
that focus on special tasks. Each sub-team has a general manager supported by several technical
leads and subordinate members. Team memberships were selected based on each individual's area
of expertise and personal interest. Error! Reference source not found. shows the work
breakdown structure of Team ARES.

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Figure 1: Team ARES Work Breakdown Structure
1.3. Launch Vehicle Summary
The Skyron Launch Vehicle is 93.67 inches in length and projected to weigh 28.00 lb. with a 30%
mass margin. Skyron is designed to accommodate a 3.5 inch PVC pipe payload in the payload bay
located just before the nose cone. A Cesaroni Technology L990 reloadable rocket motor was
chosen to propel the rocket to an apogee of 5280 ft. A 2.5 ft. diameter drogue parachute will deploy
from a compartment between the booster and avionics sections an apogee, and a 4.3 ft. diameter
main parachute will be deployed below 700 ft. AGL to slow the rocket such that the kinetic energy
at ground impact will be below 75 ft.-lbf.
1.4. AGSE Summary
Team ARES’ Autonomous Ground Support Equipment (AGSE) mission will be to secure the
payload, raise the launch vehicle, and insert the igniter. The AGSE weighs a total of 60 lbs, has a
10 ft. by 4 ft. base, and a starting height of 1.5 ft. The Robotic Payload Delivery System (RPDS),
using 4 servo motors, will deliver and secure the payload inside the payload bay of the launch

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vehicle. The Rocket Erection System (RES), actuated by a bipolar stepper motor, will raise the
launch vehicle from a horizontal position to a position 5 degrees from the vertical. The Motor
Ignition System (MIS) will use a rack and pinion system powered by a unipolar stepper motor to
insert the igniter. All the functions of the AGSE will be controlled by an Arduino Uno.

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2. Changes Made Since PDR
2.1. Launch Vehicle Changes
•
2.2. AGSE Changes
• Made base wider for increased support
• Added pulley to RES
• Changed materials in MIS from plastic to steel
• Decreased RPDS weight
• Added more ball bearings MIS
• Added electronics containment unit
2.3. Project Plan Changes
• Updated deadlines to accurately reflect the status of the project
• Update to the upcoming dates of launches and outreach efforts

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3. Launch Vehicle Criteria
3.1. Design and Verification of Launch Vehicle
3.1.1. Mission Statement
To maintain a sustainable team dedicated to the gaining of knowledge through the designing,
building, and launching of reusable launch vehicles with innovative payloads in accordance with
the NASA University Student Launch Initiative Guidelines.
3.1.2. Mission Success Criteria
Table 2: Mission Success Criteria
Requirement Design feature to satisfy that
requirement
Requirement
Verification
Success Criteria
Reach an
altitude of 5,280
ft. as accurately
as possible.
The A.T.S. will deploy during
cruise flight to adjust the flight
profile curve to match a real-
time ideal projection of the
rocket’s trajectory for the
designated altitude by increasing
the drag coefficient of the launch
vehicle.
Gathering data post-
launch from the on-
board altimeters.
The A.T.S.
directs the
launch vehicle to
an accuracy in
apogee of 2%.
The vehicle
must be
reusable.
Robust materials will be selected
for the components of the launch
vehicle that will be subjected to
high-stress environments.
By inspecting every
element of the
launch vehicle to
ensure no structure
was compromised
No visible
structural
damage is
visible and every
component is
still functional
The payload
must be retained
at all times
during flight
A payload bay with secure
payload holders will provide
sufficient force to prevent
detachment due to vibrations.
By inspecting the
payload bay post-
launch for partial or
complete
detachment.
The payload will
remain in the
same position as
it was pre-
launch.

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3.1.3. Major Milestone Schedule
The following table represents the major milestones Project Hermes will encounter throughout the
entirety of the competition. For a more detailed timeline please visit Appendix XX for a complete
view of major deadlines, specific manufacturing dates, located in the 2015-2016 Gantt Chart.
Table 3: Major Milestones
Deadline Date
Team Formation 20 AUG
Proposal 11 SEPT
Web Presence Established 23 OCT
PDR Documentation 6 NOV
PDR Teleconference 9-20 NOV
CDR Documentation 15 JAN
CDR Teleconference 19-29 JAN
AGSE, Flight Systems, and
Launch Vehicle Testing
29 JAN –
20 FEB
Full Scale Testing and
Launching
20 FEB
FRR Documentation 14 MAR
FRR Teleconference 17-30 MAR
Competition 13-16 APR
PLAR Documentation 29 APR

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3.2. System Design Review
The launch vehicle will be 93.67 inches long with a body tube diameter of 5.00 inches. The size
of the rocket was chosen to allow spacing for the parachute bays, the main avionics bay, and the
motor. OpenRocket simulation predicts that the Cesaroni L990 motor will result in an apogee of
5280 ft. with an extra mass of 615 grams to account for unexpected component additions. The
launch vehicle is divided into three sections: the booster section, avionics section, and payload
section. Table 4 lists materials used with motivation for material selections.
Table 4: Material Selection
Component(s) Material Motivation
Body Tube/Nosecone G12 Fiberglass Resistance to high aerodynamic
loads and ground impact
Fins G10 Fiberglass Resistance to high aerodynamic
loads and ground impact
Bulkheads Plywood Cheap, lightweight, and reliable

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The following Figures will provide a complete overview of Skyron’s separate components. Figure
2,3,4, & 5 represents a complete Solidworks model of the launch vehicle.
Figure 2: Overview of Skyron
Figure 3: Closeup of Fully Deployed ATS

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Figure 4: Top View of a fully deployed ATS
Figure 5: Top View of retracted ATS

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Figure 6: Close up of Fin Section
3.2.1. Final Motor Selection
The final motor choice for Skyron is a Cesaroni L990. We have selected this motor due to the
increased space capacity it provides for the Apogee Targeting System within the Booster Section.
Additionally, the L990 motor still has the necessary thrust to overshoot the target altitude of 5280
ft. so our ATS can be activated. The following Table XX outlines the specifications of the L990.
Table 5: L990 Specifications
MOTOR NAME Cesaroni L990
DIAMETER 54mm
LENGTH 64.9cm
PROP WEIGHT 1.369kg
TOTAL WEIGHT 2.236kg
AVG THRUST 991.0N
MAX THRUST 1702.7N
TOTAL IMPULSE 2771.6
BURN TIME 2.8s

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3.3. System Level Functional Requirements
Requirement Design Feature to Satisfy
Requirement
Verification
Method
Status
Vehicle altimeter will report an
apogee altitude of most nearly
5,280 feet AGL.
Low-mounted electric-controlled
fins will be extended and retracted
in reaction to altimeter readings to
control drag and limit altitude.
Analysis In
Progress
Launch vehicle will be
designed to be recoverable and
reusable within the day of
initial launch.
Vehicle will be constructed of
fiberglass to resist fractures and
ensure stability.
Design
Review
In
Progress
Vehicle will be prepared within
2 hours and will be able to
maintain launch-ready position
for at least 1 hour.
Compartmentalized design with
standard assembly procedure.
Execution In
Progress
The launch vehicle shall have a
maximum of four (4)
independent sections.
Three (3) sections include:
payload, avionics, and booster
Inspection In
Progress
The vehicle will be limited to a
single stage, solid motor
propulsion system, delivering
an impulse of no more than
5,120 Newton-seconds.
Single-staged design that utilizes
a single “L” impulse classification
motor.
Design
Review
In
Progress
Team must launch and recover
both a subscale and full scale
model prior to each CDR and
FRR respectively.
Efficient Recovery System with
redundancies to ensure successful
operation.
Execution In
Progress
The launch vehicle shall stage
the deployment of its recovery
devices, where a drogue
parachute is deployed at
apogee and a main parachute is
deployed at a much lower
altitude.
Redundant altimeters
programmed to deploy at specific
altitudes.
Inspection In
Progress

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At landing, the launch vehicle
shall have a maximum kinetic
energy of 75 ft-lbf.
Optimization of parachute sizing
for the total mass of the launch
vehicle
Testing In
Progress
The recovery system will
contain redundant altimeters,
each with their own power
supply and dedicated arming
switch located on the exterior
of the rocket airframe
Install a master key-switch at the
rear of the avionics bay to close
all circuits simultaneously, and
independent compartment for
sensors and power supply.
Inspection In
Progress
Each detachable section of the
vehicle and payload must
contain an electronic tracking
device and continue
transmission to the ground
throughout flight and landing.
Independent GPS compartment
with transmission capabilities and
ground station with receiving
capabilities.
Inspection In
Progress
3.4. Manufacturing Approach
3.4.1. Payload Section
The payload bay will be a cutout into the airframe of Skyron which will have latch to secure the
payload. In order to ensure maximum efficiency and achieve the precise dimensions for the
incisions on the fiberglass airframe, a conventional CNC Mill was selected for this manufacturing
process. This choice also guarantees the least amount of glass fibers delaminating from the epoxy
composite, and under the correct safety procedures only produces fiberglass residues in the very
manageable form of dust, minimizing the risk of toxic fume generation. This process will be
utilized to manufacture the hatch for the payload bay which should maintain its structural integrity
for its exposure to the high velocity flow during ascent.

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In order to ensure maximum efficiency and achieve the precise dimensions for the incisions on the
fiberglass airframe, a conventional CNC Mill was selected for this manufacturing process. This
choice also guarantees the least amount of glass fibers delaminating from the epoxy composite,
and under the correct safety procedures only produces fiberglass residues in the very manageable
form of dust, minimizing the risk of toxic fume generation. This process will be utilized to
manufacture the hatch for the payload bay which should maintain its structural integrity for its
exposure to the high velocity flow during ascent.
To house the payload bay in the upper section, two plywood bulkheads enclose the bay on either
end.

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The procedure for cutting plywood of various thicknesses involves using a high powered laser
cutter. In order to create a component with exact dimensions, a two dimensional computer-aided
design (CAD) file is drawn and then input through Inkscape into the high powered laser cutter.
This procedure is fast and effective with an accuracy of a thousandth of an inch, allowing for near
perfect assemblies when it comes to multiple components. For example, this method was utilized
for the upper section to create the various thickness bulkheads as well as the support structure
within the payload bay. This method guarantees that the manufacturing process can occur without
any hazards to safety.
Another method for manufacturing to be utilized for the upper section of the launch vehicle is
additive manufacturing. This method provides a time effective method of transferring a three
dimensional CAD file into solid structural material with the ability of having high reproducibility.
The material with which the 3D printers operate is ABS plastic, which provides enough structural
integrity and ease of handling to be used for more complex and delicate components of the design.
It is for these reasons that additive manufacturing was utilized to create the payload holding clips
that allow for slight deformation with full elastic recovery for the process of inserting and
removing the payload.
This material is also lightweight enough to be non-intrusive to the overall design of the launch
vehicle so this method also allows for high margins for the modification of such components.

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3.4.2. Avionics Section
3.4.2.1. Main Avionics Bay
The avionics bay consists of two different materials: G12 fiberglass and 0.125in plywood boards.
These two materials require different manufacturing methods to ensure that their structural
integrity isn’t permanently affected. For the plywood boards, the conventional method for altering
the dimensions of the board is using a high powered laser cutter for precise and safe manufacturing.
As to the fiberglass tubes, what is most convenient is to use a table saw, while still taking into
account the safety hazards that arise from cutting fiberglass, so the appropriate safety equipment
must be used by every individual present during the time of manufacturing. As to the holes that
secure the Avionics Bay in place, a conventional drill will be used while still accounting for the
same safety hazards as previously discussed. These methods ensure there will be little deformation,
delamination, and precise cuts for the manufacture of each component.
The Avionics bay will house the components in charge with the recovery system, ATS system,
and data collection system. Skyron’s Avionics Bay (AB) is where all the board readings,
measurements and information is processed. To house all of the necessary avionics components

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the AB will be placed on a 10” x 4.8” x 1⁄8” vertical plywood board which is a epoxied onto the
smaller diameter of the inside of the avionics bay.
The avionics bay itself consists of two symmetrical sections one of which is attached to the main
body of the rocket and the other which can be easily inserted and removed via connection pins
placed on the epoxied vertical board and metallic tabs extending from the edges of the inner
diameter of the removable component. The key switch for the ignition will be placed on the
avionics bay security latch which is screwed in with two standard ⅛” screws. The two security
latches/doors are screwed in on both sides of the avionics bay with access to either side of the
epoxied vertical board. Some major setbacks were found with the previous design in the subscale
launch insertion, one of which was the difficulty of inserting the avionics “sled” in the avionics
bay due to space congestion generated by the large number of wires and avionics equipment. This
resulted in the “sled” not coinciding flush with the door cutout. Through the aforementioned new
design, these insertion and setup difficulties are eliminated leading to ease of access, installation
and reinforced security for the avionics components and bay in its entirety.
3.4.2.2. Nosecone GPS
The GPS will be located within the nosecone of the launch vehicle to ensure that the signals do
not interfere with the rest of the on-board electronics. To secure it in place, it will be mounted on
a construction of ⅛” thickness plywood and with its base on the payload bay. The GPS will be
mounted on one side of the plywood board while its isolated power supply is located on the other
side of the board. This construction only requires a high powered laser cutter to manufacture the
boards with the right dimensions and the right holes for the attachment points of the electronics.
Attaching the GPS Bay to the rest of the launch vehicle will require it to be epoxied to the top of
the bulkhead enclosing the payload bay, allowing it to be easily inserted into the nosecone before
securing the latter in place with the assistance of screws.

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3.4.3. Booster Section
The Booster Section (BS) of Skyron will house the L990 Motor, and the Apogee Targeting System
(ATS). For more information concerning the Motor selection please see Section XX. The Motor
Retention Plate, which will hold the motor in place and prevent the motor from travelling straight
through the Launch Vehicle, will be manufactured from ½ “ plywood. To ensure that the maximum
thrust of the motor does not penetrate and create enough displacement to cause problems, we
simulated the stress on the motor retention plate using Finite Element Analysis. Finite Element
Analysis (FEA) is a numerical technique used for finding approximate solutions to partial
differential equations. This technique is useful for theoretical analysis of design components.
Solidworks utilizes this technique to perform basic FEA, and was used to analyze the thrust plate,
shown below
The force applied corresponds to the maximum thrust the Cesaroni L990 can produce, which is
382.62 lbf. The maximum displacement of the thrust plate was 0.01 inches with a maximum stress
of 375 N/m^2. The figures were scaled to emphasize the displacement.

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Another main component of the BS is the Apogee Targeting System (ATS). To assemble the ATS,
four 3-D printed mounting frames will be evenly spaced around the booster.
The frame consists of custom fit cut-outs for the ATS motor and battery to be placed inside. The
bolt and lead screw will slide into the mounting frame and connected to the DC motor. The ATS
tab will be attached to a hinge inside the body tube wall. The pivoting rods, which translate the
bolts’ linear motion to the tabs, will have thin cut slots in the body tube to allow motion both inside
and outside of the body tube.
The ATS will be powered by four (4) DC motors. Each of the four ATS tabs will be connected to
a lead screw powered by its own DC motor and battery incorporated within each separate mounting
frame.
The fins will be manufactured out of G10 Fiberglass sheets using a Maxiem 1515 CNC Waterjet.
This method has proven to be successful in the past leading to an accurate cut with little to no
safety concerns.

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3.4.4. Status and Plans of Remaining Manufacturing and Assembly
As of CDR, every major component is being assembled; the booster section, the avionics section,
and the nosecone section. As more and more components arrive, more manufacturing will take
place. The structural integrity of every part is being considered as we move toward the full scale
testing.
3.4.5. Integrity of Design
3.4.5.1. Suitability of Shape and Fin Style
The fins will be made using G10 Fiberglass as the material of choice. Initially, the fins were
attempted to be made with a smooth airfoil shape in order to improve the aerodynamics of the fin
and reduce drag. Due to complications in the sanding process, it was determined that the smooth
airfoil shape would be unreasonable for the fins due to the fact that G10 Fiberglass is not one solid
material, but multiple layers on top of each other. During sanding, it is expected that the layers
would begin to peel off one another.

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The fin has a clipped delta fin shape (as shown below) which was determined as the most viable
option for a launch vehicle with four fins. With four fins, the stability of Skyron will increase as
opposed to using only three fins (stability is expected increase by slightly over 50%). The fin flutter
speed was calculated using the Flutter Boundary Equation published in NACA Technical Paper
4197:
The corresponding variables for our fin are listed in Table 12 located below. The fin flutter speed
was calculated to be 1326.109 mph. Comparing Vf to our maximum velocity Vmax of 552.148 mph
(0.72 Mach), Skyron will not experience the unstable effects of fin flutter. Exceeding the fin flutter
speed will exponentially amplify the oscillations and rapidly increase the energy in the fins;
causing greater induced moments and more instability.
Table XX: Fin Dimensions
Variable Unit
Speed of Sound, a 1105.26 ft/s
Pressure, P 13.19 lbm/in2
Temperature, T 48.32 Fahrenheit
Shear Modulus, G 425,000 psi
Taper Ratio, 0.3627
Tip Chord 7 cm or 2.75591 in
Root Chord 19.3 cm or 7.598 in
Thickness 0.318 cm or 0.1252 in
Fin Area 55.23 in2
Span 13.4 cm or 5.275591 in
Aspect Ratio 0.50392

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3.4.5.2. Material Analysis
The fins of the rocket are going to be made of G10 fiberglass. This fiberglass is being used
for its high ultimate breaking point and its ductility, which allows the fins to have more flexibility.
This means the fins are more likely to stay intact during a soft. and hard landing without cracking
as compared to previously used materials like carbon fiber and wood. However, wood is being
used for the bulkheads, specifically ¼ inch plywood. Plywood is much cheaper than fiberglass and
it can meet expected standards of performance at maximum thrust. The use of plywood has also
been proven to be more efficient in manufacturing the bulkheads using the machinery we have at
hand, such as the laser cutter, since we have already cut bulkheads using plywood for the subscale
launch in relatively short time. The materials being used for the the body tubes and the nose cone
is G12 fiberglass. This means our team has the ability to have more test launches with the same
materials since the overall exterior of the rocket has high durability, thus G12 fiberglass is the best
choice for our final launch.

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3.4.5.3. Motor Mounting and Retention
Motor mounting shall be undertaken by our mentor and NAR Level II Member Joseph Mattingly.
No other personnel on the team is allowed to handle the motor as it would be a huge safety risk
and Joseph Mattingly is the only person qualified within our team to build and mount a rocket
motor. To view our retention system and analysis, please refer to section 3.4.3.
3.4.5.4. Mass Statement
The mass of the launch vehicle is depicted in the figure below. The different categories are defined
by what purpose they serve in the launch vehicle’s performance. Combined, all the components of
the vehicle have a total mass of 9747.6 grams. This is an educated estimation of what the total
mass of the rocket will vary. Of course, this mass estimation isn’t absolute, since a growth of 25-
33% was accounted for since the submission of the Project Proposal.
Currently, a more complete Mass Statement with the weight of every manufactured and purchased
component is being undertaken.

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3.5. Subscale Flight Results
3.5.1. Flight Data
Found below is our OpenRocket Simulation that allowed us to design a subscale with similar
specifications and design features.
The figures below demonstrate the successful launch and recovery of the subscale launch vehicle.
The first figure depicts the constant voltage supply from the 9-volt batteries to the altimeters, with
the exception at the times in which current was diverted onto the blasting charges for the
deployment of each parachute. Cross-checking the deployment times with the second figure, it is
visible that both deployments occurred at the right altitudes recorded by the Stratologgers. The
subscale launch vehicle attained a maximum altitude of 2113 ft., with the assistance of fully
deployed scaled models of the ATS tab designs. The additional drag generated by these tabs
disrupting the flow all throughout the flight reduced the expected apogee significantly.

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3.5.2. Result Discussion
The figure below demonstrates an apogee 2506 ft., which is 393 ft. higher than the experimental
results with full deployment of the subscale ATS. This is a 15% reduction in altitude generated
from the additional drag generated from the disturbance of the laminar flow around the body of
the launch vehicle, sufficient enough to compensate for any unexpected variations in atmospheric
conditions, mass margin errors, and other unpredictable phenomena. The experimental flight also
provided us feedback on aspects that need to be deemed of high importance to guarantee the safe
operation of the launch vehicle such as the audible feedback from the altimeters, battery life,
accessibility of all the switches, and the ease of separation of the multiple segments of the launch
vehicle.

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3.5.3. Impact on Full Scale
Throughout the manufacturing and assembling process of the subscale, many of the designs
resulted impractical and mechanically more complex than expected, thus major redesigns had to
occur in order to guarantee the functionality of all subsystems. The avionics bay wiring created
many issues for assembling the subscale due to the lack of available space within the avionics bay,
and thus cable management will be one of the priorities for the assembling procedures of this
section of the launch vehicle. Other changes were the position of the GPS, the power supplies, and
the key switches, since many of this caused a complication for the time of assembly and pre-launch
checks. Another area that was subject to change was the fin assembly method, since the
impracticality of having fin braces was not outweighed by its benefit, and so the attachment points
for the fins were increased in surface area contact which also reduces the overall weight of the
booster section. The positioning of the payload bay was also reworked, due to the volume
constraints for the parachute packing and preparation for launch; thus more space will be allocated
to enclosing both parachutes along with their respective shockcord and insulator.
Pre-flight checklists will also be essential to the success of the full scale test launch, since the
assembly of such vehicle requires numerous steps. Pre-checking every subsystem and every

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electronic system must be done with plenty of anticipation to ensure that everything is working
perfectly and to avoid any last minute inconveniences at the launch pad.
Most importantly, the aerodynamic effect of the ATS proved itself to be significant enough to
justify the complexity of the design, even with only two deployed tabs rather than the full 4-tab
configuration. The mechanical component of these tabs was reworked due to the confined space
and mass limitations encountered during subscale manufacturing. The structural soundness of the
airframe was also compromised within safe boundaries, but appropriate design changes were
accommodated to minimize the added risk of multiple incisions on the structure.

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3.6. Recovery Subsystem
3.6.1. Recovery System Hardware
The main parachute of diameter 50” will be housed in the avionics section, while the drogue
parachute will be located just below the avionics section (using the nosecone as a reference
location). Both parachutes will be fabricated from rip-stop nylon in order to support the weight of
the launch vehicle. Parachutes will be secured in their individual sections using an insulated
material to prevent the ignition of the nylon due to explosive charges that will separate the different
sections of the rocket sections during descent deployed from the blasting caps that are attached to
4.85” diameter bulkheads which seal the avionics bay from the rest of the rocket’s compartments.
The parachutes will be attached and secured to the rocket via the shock-cords which are connected
to U-bolts installed onto the respective bulkheads/centering rings insulating each section of the
rocket from pressurization.
Parachute Diameter (inches)
Main 50
Drogue 15
3.6.2. Electrical Hardware
The StratologgerCF altimeter records data at a rate of 20 samples per second and stores it for later
use. They also include a Data I/O connector which allows for real-time altimeter data to be sent to

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the onboard flight computer. This altimeter is functional up to an altitude of 100,000 feet and will
be used to deploy the main and drogue parachutes upon reaching specified altitude.
3.6.3. Kinetic Energy Calculations
The total mass of the launch vehicle is given as 9747.6 grams. The nose cone will be made
fiberglass and has a weight of 214.74g or 0.47lb. The rocket’s propulsion system/booster section
accounts for 22.94% of the overall mass of the launch vehicles as shown in the chart below. The
weight of the booster section after the propellant has been expended was estimated at 1.5 lbs
The equations below were used to estimate the ideal impact velocity, v of the launch vehicle
assuming no external forces.
KE = ½ mv2
Where v is the ground hit velocity of the launch vehicle, 10.6m/s or 34.77ft/s.
Sections Dry Mass (lbs) KE (ft-lbf)
Nosecone 0.47 8.83
Avionics 2.2 35.44
Booster 1.5 28.18
Total 72.45

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3.7. Mission Performance Predictions
3.7.1. Validity of Analysis
Due to the design compromises made during the subscale launch, the validation of the drag
coefficient and effect of the ATS on the rocket could not be conducted. We did acquire promising
data to the subsections usage and will further pursue experimentation with a new deployment
system and placement on the rocket. The ATS has been placed 18 inches from the bottom of the
rocket to accommodate the motor driven system. With the new position, we must investigate any
effects the turbulent flow will possible have on the stability of the rocket
To illustrate the effect of the ATS on the rocket, we have used CFD analysis from Ansys Fluent
16.2. An attempt was made to use the Flow Simulation add in of Solidworks 2015 for convenience
and validation, but the results were highly unreliable. Ansys provided us with answers fairly
similar to the original CFD files. When we take a look at the effects of the turbulent kinetic energy,
it’s possible to see how raising the flaps of the ATS by approximately 10 inches has created a wake
of turbulent flow interrupting the laminar flow to the fins. This increases our overall drag by
increasing pressure drag. The pressure contour below shows the mixing of the low and high
pressures around the flaps creating our turbulent flow.

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3.7.2. Drag Assessment

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3.8. Future Testing
Component Test Verification
Method
Lead Screw with DC
motor actuation
Extension force of flaps test. Quantitative
Analysis
ATS Wind tunnel testing to confirm Cd simulations. Quantitative
Analysis
Thrust Plate Bend test and pressure test to verify rigidity until
breaking point.
Quantitative
Analysis
Payload Bay Payload retention force measurement test. Quantitative
Analysis
Avionics Bay Altimeter accuracy and accelerometer performance
test.
Quantitative
Analysis
Recovery System Recovery system ground test fire. Inspection
Fins Fin attachment robustness test along two axis. Quantitative
Analysis
Launch Vehicle
Assembly
Vehicle will be completely assembled under a time
constraint to verify efficiency and effectiveness.
Inspection

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3.9. Launch Procedures
3.9.1. Launch Checklist
Below is a preliminary checklist of all the procedures and steps to follow to have a successful and
safe launch.
Prepare Payload Recovery System
Ensure batteries and switches are wired correctly
Ensure batteries, power supply, switches, microprocessor, GPS, XBee is/are wired correctly
Install and secure fresh batteries into battery holders
Insert payload recovery electronics into payload recovery bay
Connect appropriate wires
Arm altimeter with output shorted to verify jumper settings. This is done to verify battery power and continuity
Disarm Altimeter, un-short outputs
Insert Payload Recovery Bay into Payload Section
Prepare Body Recovery System
Ensure batteries and switches are wired correctly
Ensure batteries, power supply, switches, microprocessor, GPS, XBee is/are wired correctly
Install and secure fresh batteries into battery holders
Insert body recovery electronics into payload recovery bay
Connect appropriate wires
Arm altimeter with output shorted to verify jumper settings. This is done to verify battery power and continuity
Disarm Altimeter, un-short outputs
Insert Body Recovery Bay into Payload Section
Assemble Charges
Test e-match resistance to see if it is within specifications
Remove protective cover from e-match

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Measure amount of black powder used in testing
Place e-match on tape with sticky side up
Pour black powder over e-match
Seal Tape
Re-test e-match
Check Altimeters
Ensure altimeters are disarmed
Connect charges to ejection wells
Turn on altimeters to verify continuity
Disarm altimeters
Pack Parachutes
Connect drogue shock cord to booster section and body section
Attach drogue parachute to drogue shock cord
Pack drogue parachute
Fold excess shock cord so it does not tangle
Attach Nomex cloth to shock cord so it will enclose and shield the parachute while exposing only the Kevlar shock cord to
ejection charge
Insert cellulose wadding into drogue parachute bay between ejection charges and parachute
Insert drogue parachute and shock cord into drogue parachute bay
Insert booster section into lower body section, and secure with shear pins
Attach main parachute shock cord to upper body section and lower payload parachute bay
Attach main parachute to main parachute shock cord
Pack main parachute
ejection charge
Insert cellulose wadding into main parachute bay between ejection charges and parachute

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Insert main parachute and shock cord into main parachute bay and
Insert upper body section into the lower section of the payload parachute bay, and secure with shear pins
Attach payload parachute shock cord to payload section
Attach parachute to the end of the payload parachute shock cord
Pack payload section parachute
ejection charge
Insert cellulose wadding into upper payload parachute bay between ejection charges and parachute
Insert drogue parachute and shock cord into upper payload parachute bay
Insert payload section into payload parachute bay and secure with shear pins
Assemble motor
Follow manufacturer’s instructions
Do not get grease on propellant grains or delay grain
Do not install igniter
Install Motor in launch vehicle
Secure motor retention system
Launch Vehicle Prep
Inspect launch vehicle, check CG and make sure it is within specified range
Bring launch vehicle to Range Safety Officer(RSO) for inspection
Bring launch vehicle to Autonomous Ground Support Equipment(AGSE) platform
Install launch vehicle on AGSE
Install motor igniter on AGSE
Touch igniter clips together to make sure they will not fire the igniter when connected
Make sure igniter clips are not shorted to each other or any section of the AGSE
Connect igniter clips to motor igniter

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AGSE Prep
Activate AGSE master switch and ensure safety light is flashing in color
Activate AGSE pause switch and ensure safety light is solid in color
All nonessential personnel evacuate to safe launch distance
Deactivate AGSE pause switch and start stopwatch to time AGSE routines
Stop stopwatch when AGSE routines are complete and record time from pause switch deactivation to rocket erection
Essential personnel will arm altimeters via switches and ensure continuity
All personnel will evacuate to safe launch distance
Launch
Watch flight so launch vehicle sections do not get lost
Post Launch Payload/Vehicle Recovery
Recover Payload Section and tethered Body/Booster Section
Disarm Altimeters if there are unfired charges
Disassemble launch vehicle, clean motor case, other parts, and inspect for damage
Record altimeter data

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3.10. Safety and Environment
3.10.1. Overview
Team A.R.E.S. is dedicated to maintaining safe operating conditions for all team members and
anyone involved in competition activities. Under the tutelage of the Safety Officer, Team A.R.E.S.
will undergo rigorous safety briefings to ensure the integrity and safety of the entire team and
equipment is unchanged. During manufacturing, fabrication, and testing of rocket vehicle and
AGSE components, it is important to identify the hazards of your environment, and how following
safety procedures and protocols can prevent accident and injury to oneself or damage to
competition hardware. When working with construction equipment, Team A.R.E.S. members are
instructed to work in minimum team sizes of two. This ensures that one team member would be
available to provide immediate assistance or quickly get help should an incident occur while using
the equipment. The Invention Studio, where team members use the necessary equipment for
manufacturing and fabrication, is equipped with first aid kits, fire extinguishers, safety glasses,
and expert supervision for the use of all equipment. During physical testing of the rocket structure,
and during ejection charge testing, team members will wear safety glasses, have a first aid kit and
fire extinguisher on hand, and have licensed safety officials present. In order to use the machines,
all team members have been briefed on the proper protocols and procedures of using the lab
machines. Risk identification and mitigation techniques are used to assess the dangers of tools and
activities to personnel, and how they may create safe operating conditions. To that end, Table XX
lists the procedure to identify what hazards and risks may exist and how to minimize the chances
of occurrence.
Step Name Step Definition
1. Hazard
Identification
Team will collectively brainstorm to identify any possible
hazards that the team may encounter.
2. Risk and Hazard
Assessment
Team will determine the severity and probability of
consequences in case the hazard were to be encountered. How to
approach each hazard will also be reviewed.
3. Risk Control and
Elimination
After the hazard has been identified and assessed, a plan
will be put in place to ensure the hazard will not occur.
4. Reviewing
Assessments
The entire process will be repeated for any new hazards
or existing hazard that needs to be updated.

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3.10.2. Failure Modes and Analysis
The following table entails the information on the general risks that the team may experience
while constructing the rocket. The table goes through the severity, likelihood, and prevention
methods for each hazard.
Hazard Severity Likelihood Mitigation & Control
Batteries
Explode
Burns, skin and
eye irritation
Low Wear safety glasses and gloves
when handling. Make sure no
shorts exist in circuits using
batteries. If battery gets too hot,
stop its use and disconnect it from
any circuits.
Black Powder Explosions,
burns, skin and
eye irritation
Medium Wear safety glasses, gloves when
handling black powder. Be careful
when pouring black powder.
Operate in a static-free
environment
Dremel Cuts and scrapes Medium Only operate tools with supervision
of teammates. Use tools in
appropriate manner. Wear safety
glasses to prevent debris from
getting into eyes.
Power Tools Cuts, punctures,
and scrapes
Medium Only operate power tools with
supervision of teammates. Use
tools in appropriate manner. Wear
safety glasses to prevent debris
from getting into eyes.
Epoxy/Glue Toxic fumes,
skin and eye
irritation
High Wear gloves, nitrile for epoxy, face
masks, and safety glasses. Work in
well ventilated area.
Exacto/Craft
Knives
Cuts,
serious/fatal
injury
Medium Only use knives with teammate
supervision. Only use tools in
appropriate manner. Do not cut in
the direction towards oneself.

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Fire Burns,
serious/fatal
injury
Low Keep a fire extinguisher nearby. If
an object becomes too hot, or does
start a fire, remove power (if
applicable) and be prepared to use
the fire extinguisher.
Hammers Bruises,
serious/fatal
injury
Medium Be aware of where you are
swinging the hammer, so that it
does not hit yourself, others, or
could bounce and hit someone.
Hand Saws Cuts,
serious/fatal
injury
Medium Only use saws with teammate
supervision. Only use tools in
appropriate manner. Wear safety
glasses to prevent debris from
getting in eyes.
Waterjet Cutter Cuts,
serious/fatal
injury, flying
debris
Low Only operate under supervision of
Undergraduate/Graduate Learning
Instructors, and with other
teammates. Follow proper
operating procedures, wear safety
glasses.
Improper dress
during
construction
Cuts,
serious/fatal
injury
High Wear closed toed shoes, tie back
long hair, do not wear baggy
clothing.
Power Supply Electrocution,
serious/fatal
injury
Medium Only operate power supply with
teammate supervision. Turn off
power supply when working with
circuitry.

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The following table entails the potential failure modes that may be experienced by the Launch
Vehicle team and the prevention method for each failure mode.
Potential
Failure
Effects of Failure Failure Prevention
Apogee
Targeting
System
(ATS)
Vehicle will not reach target
altitude
Test ATS using subscale launch vehicles
Body
structure
buckling on
takeoff
Launch failure, damage to launch
vehicle, unable to be reused,
flying shrapnel towards
personnel/crown
Test structure to withstand expected
forces at launch with a factor of safety.
Have proper sized couplers connecting
sections.
Drogue
separation
Main parachute will deploy at
high speed and may rip or
disconnect from vehicle, launch
vehicle may become ballistic
Perform ground test and flight test.
Fins Fins could fall off, causing
unstable flight.
Fins break or disconnect from
launch vehicle, unable to be
classified as reusable
Test fin at attachment points using
expected forces to ensure strength of
attachment method.
Do not have fins with sharp pointed
edges, ensure parachute is large enough
to minimize impact kinetic energy, test
fin at attachment points using expected
forces to ensure strength of attachment.
Ignition
failure
Failure to launch Follow proper procedures when attaching
igniter to AGSE.
Launch
buttons
Launch vehicle will separate from
rail, causing an unstable flight
Ensure launch rail is of proper size to
accommodate the buttons, ensure buttons
slide easily into rail.
Main
parachute
separation
High impact velocity may damage
vehicle and make it
unrecoverable, vehicle may
become ballistic causing serious
injury or death
Perform ground test and flight test to
ensure veracity of deployment method.
Motor failure Motor explodes, damaging launch
vehicle/AGSE beyond repair
Follow NAR regulations and
manufacturer’s instructions when

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assembling motor. Assemble motor under
supervision.
Motor
retention
Motor casing falls out, lost motor
case, could damage
persons/property
Test reliability of motor retention system
Payload
separation
Main parachute may not deploy
correctly, higher impact velocity
may damage launch vehicle, or
cause personal/property damage
Perform ground and flight test to ensure
veracity of deployment method
Thrust plate
failure
Motor goes through vehicle,
damage to vehicle, causing it to
be not reusable
Test plate and attachment method to
withstand expected launch forces with a
factor of safety
The following table entails the potential failure modes that may be experienced by the AGSE
subteam and the prevention method for each failure mode.
Potential Failure Effects of Failure Failure Prevention
Payload is not
secured in bay
Payload will bounce inside
payload bay, disrupting
flight
Test various plastic clip dimensions to
find best fit
RES is not stable
while raising
Rocket will not be raised,
and potentially the motors
will be broken
Test subsystem, add counterweights to
reduce necessary force from motor, and
add more framing to increase stability
RES does not stay
upright
Launch vehicle will fall
unpredictably
Perfect ratchet system, ensure tension in
steel cable
Electronics short
circuit or are
overloaded
System will lose control Fuses will protect electronics, subscale
testing will prevent short circuits and
overloads

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3.10.3. Personnel Hazards
Personal injuries can occur at any given time throughout the entire project. Each individual should
be aware and alert at all times when working on the rocket. Warning labels on hazardous materials
should be thoroughly read. Equipment should only be used with authorized personnel present.
Each team member should be familiar with the safety hazards and prevention methods listed in
this document as well as in the safety handbook provided by NASA. Every team member
understands that the safety guidelines and procedures outlined must be followed at all times.
Failure to do so may result in injury and/or death.
3.10.4. Environmental Concerns
The team understands that building a rocket requires the use of many equipment and/or materials
throughout the entire design process. Despite the complexity of building a rocket, the environment
must be taken into account at all times. Hazardous materials must be properly disposed of.
Launches may only take place on authorized days and times. The Material Safety Data Sheet
(MSDS) for each material used must be thoroughly read by each team member. Team ARES will
do its best to ensure that the negative impact on the environment is at a minimum while designing
the launch vehicle.

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4. Flight Systems
4.1. Overview
The onboard Flight Software was designed to be modular in order to maximize performance while
also making it easy for the team to work on it cooperatively. A Scheduler function, which can be
reviewed in Appendix XX, determines the frequency at which each block of code runs in reference
to the other blocks.
The process of scheduling the blocks of code begins with polling the sensors. This code block
outputs an altitude and acceleration vector which is then input into the State Estimation block. The
State Estimation block runs five times for every single time the sensors are polled. This process
allows the State Estimation block to compare the sensed data to the ideal path of the rocket. This
process continues 20 times before the controller code block is scheduled. Based on the estimated
state, the Controller outputs whether to actuate or not. The Controller, and thus the Actuator, block
runs every 100th
time the state is estimated and every 20th
time the sensors are polled
The scheduling process allows the team the opportunity to choose how often the ATS is triggered
as well as how much data is gathered before adjusting flight path. The altitude and acceleration
data gathered during the sensor polling code block ultimately drive the decisions made throughout
the rest of the code and lead to the decision of whether to correct flight path or not.

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4.2. Simulink Design Overview
This is the general form of our simulink model. This Simulink model aim to help us first Simulate
the program and help us to design the motor we want and need. Also, this simulink allow us to
design the controller for this system.
Here is the General Simulink model we have:
We can see the controller and the system block.
In this section, we are going to present the Rocket block. We have implemented a Simulink model
for the equations of the rocket. Briefly, we have: weight, engine thrust, plates drag, and rocket drag
acting on the rocket. So, we created these forces, summed them (with the adequate projection) and
then thanks to Newton’s second law we know that this net sum equals the mass of the rocket times
its acceleration.

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In this diagram, we modeled the engine by computing a function depending only on the time. From
this function we are then able to compute the mass ejected by the motor since launch and the thrust
level. Please find the Simulink model of the motor below:
4.3. Controller Design Overview
4.3.1. Control Technique
The objective of the control here is to reach the exact apogee using flaps. Indeed, the flaps are
going to be extended depending on the relative position of the rocket and a “nominal” trajectory
pre generated by simulation. The nominal trajectory generated is generated using the equation of
the motion of the rocket: then we know that in the absence of perturbation it is exactly the motion
that the rocket should have and so we know that this trajectory is perfectly doable by the system,
Please see below an example of a nominal trajectory used for the subscale launch.

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Once this nominal trajectory generated, we are going to implement a controller such that the error
between the current altitude and the nominal altitude goes to zero as t goes to infinity.
For that, we will use here a system of flaps. The control authority we have here is therefore: either
extend the flaps and create more drag, either not extend the flaps and create less drag. With this
technique the rocket will then be able to reach an apogee that is close to the one we want even with
the presence of perturbation.
4.3.2. Sliding Mode Control
Let u be the area of the flaps we want. This will be our control input. After deriving the equations
of motion, we found that implementing a Sliding Mode controller applies very well at the
equations. Therefore, we will then use a Sliding Mode Control to control the rocket.
The control law we used is then:

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With the following:
Where :
x1 = relative altitude of the rocket from nominal trajectory
x2 = relative vertical speed of the rocket from nominal vertical speed
hp = nominal vertical speed
Crocket = drag coefficient of the rocket body
Sr = Surface of the rocket that drags
Cflaps = drag coefficient of the flaps
Spo = Surface of the flaps half-extended
ρ = density of the air
For more information, please see Appendix XX, written by a Graduate student member.

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5. AGSE Criteria
5.1. AGSE Summary
Team ARES’ Autonomous Ground Support Equipment (AGSE) mission will be to secure the
payload, raise the launch vehicle, and insert the igniter. The AGSE weighs a total of 60 lbs, has a
10 ft. by 4 ft. base, and a starting height of 1.5 ft. The Robotic Payload Delivery System (RPDS),
using 4 servo motors, will deliver and secure the payload inside the payload bay of the launch
vehicle. The Rocket Erection System (RES), actuated by a bipolar stepper motor, will raise the
launch vehicle from a horizontal position to a position 5 degrees from the vertical. The Motor
Ignition System (MIS) will use a rack and pinion system powered by a unipolar stepper motor to
insert the igniter. All the functions of the AGSE will be controlled by an Arduino Uno.

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5.2. Design Review of AGSE Equipment
5.2.1. Robotic Payload Delivery System (RPDS)
5.2.1.1. Drawings and Specifications

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5.2.1.2. Analysis Results
The Rapid Payload Delivery System was designed to be extremely simple, at a very low budget,
and yet efficient in its movements. Inserting the payload into the payload bay in the launch vehicle
requires precision, however, it does not require a significant amount of mechanical tolerance for

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failure, as the payload weight is 5 oz. To follow these design features, the robotic arm’s servo
motor mounts will be 3D printed in ABS plastic and the struts will be laser cutted out of plywood.
Inside the launch vehicle, there is a simple locking mechanical component (shown in the previous
section) to ensure there is no payload movement during flight. The locking mechanism will be 3D
printed and its plastic material properties allow a secure lock after deformation. Although, the task
of the RPDS is relatively simple, it requires precision. As a result, infrared sensors will be
implemented into mechanical structure of the claw of the robotic arm, so as to locate the payload
and its delivery position.
5.2.1.3. Integrity of Design
As explained above, the key of the RPDS design is simplicity. However, various problems may
arise from adopting a simple mechanical system. One aspect to consider wisely is the materials
used. During the manufacturing and assembly process of the robotic arm, problems with the
strength of the 3D printed servo mounts were encountered. ABS plastic proved to be relatively
brittle, and thus forced the design of the mount to change and account for the stresses applied to
the part. If the 3D printed components continue to fail mechanically, these components will be
manufactured through milling operations on a delrin stock. Again, the RPDS was design to be
minimalistic and cheap. However, if this design encounters more problems with the manufacturing
and assembly process, it shall be forced to change dramatically. This change will result in the
purchase and use of stronger servo motors or robotic arm kits.

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5.2.2. Rocket Erector System (RES)

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The launch vehicle is raised from a horizontal position to 5 degrees off of the vertical position
using the force of a stepper motor winding a spool of cable around a pulley system. A cable/pulley
system is an easy way to manipulate the direction vector of force, and allow for lifting the Launch
Vehicle in a controlled manner. The stepper motor provides enough torque to rotate the spool to
wind the cable, thus raising the launch vehicle. The stepper motor is fastened to the frame using
T-nuts. The spool is attached to the stepper motor using a force fitted coupler. The coupler is
secured inside the spool with screws. The cable is fastened to the spool using a tightened hose
clamp. Then it is strung upwards around a pulley, which is also connected to the vertical part of
the frame using T-nuts, and then tied and clamped to a fixed eye screw on the lifting rod across
the length of the frame. As the stepper motor rotates the spool, the cable is wound around the spool
causing tension to pull the lifting rod and raise the launch vehicle.
The stepper motor winding the cable around a spool to create pulling tension on the lifting rod is
strong enough to lift. the Launch Vehicle from a horizontal position to 5 degrees off of the vertical
position. The stepper motor provides ample torque, and the steel cable can withstand a much larger
force than the amount needed to lift. the launch vehicle. However, a taut cable is not extremely
stable, and the launch vehicle or cable may rock during raising. The frame has a wide base to help
fight potential instability. After determining the cable’s stability in testing, a cable guide may be
added to the design if needed. Having the cable be strung upwards around the single pulley and
tied to the lifting rod on the other end of the frame puts a lot of of stress on the pulley and the T-
nuts it is fastened to the frame with. If the pulley appears to be under too much stress during testing,
a second pulley will be added to the system to relieve tension from the original single pulley. The
cable will then be strung upwards and interweaved between the two pulleys before it reaches the
lifting rod.

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5.2.3. MIS

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This rack and pinion system was chosen for its reliability. With both the guide pieces and ball
bearings, the rack is accurately moved into the motor cavity. The guides and the bearings keep the
rack steady as the pinion rotates. The steel used to construct the subsystem also helps act as
counterweight when raising the launch vehicle. The guide rails are held in place by the steel side
pieces which are fastened to the end of the main rail. The rack can extend to 1.5 ft. past its starting
position.

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The design’s strength comes from its ease of construction. The steel guides and sides can be
manufactured using a water jet. The water jet ensures greater accuracy than many other machining
methods during construction. The pieces can be quickly assembled by fitting the guides into the
sides, then the sides to the rail using standard fasteners. The steel sides have multiple holes for
bolts so they will be stable along the rail. More guides may have to be added if the rack is not
stable enough during motion. Another possibility is that the guides contribute too much friction.
Then, a new guide design may have to be created. Future redesign may also be necessary if a
coupler is needed for the stepper motor.

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5.2.4. Electronics

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The unipolar stepper motor is used to raise the Launch Vehicle, and the bipolar stepper motor is
used to power the MIS. The 5 servo motors comprise of the 4 degrees of rotation of the arm and
the claw. Each major component of the AGSE has its own green and red LED to indicate whether
that specific task has been accomplished. The yellow LED only turns on when the entire system is
in the pause state. The emergency kill switch will be connected between the battery and the
Arduino, so it directly cuts the power to the entire system. An Arduino Mega was chosen because
it contains enough digital inputs and outputs while also providing the necessary 5V of operating
voltage. It was also the cheapest option of all the microcontrollers that met those guidelines.
Component Quantity
Unipolar Stepper Motor 1
Bipolar Stepper Motor 1
Servo Motors (Tower Pro MG 995) 5
LEDs 7 (3 Red, 3 Green, 1 Yellow)
Arduino Mega 2560 - R3 1
Stepper Motor Drivers - (ROB-12779) 2
Emergency Kill Switch 1
Start/Pause Button 1
Electronics Box 1
Resistors and Wires As Needed

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The design for the RPDS can accurately capture, secure, and deliver the payload because of the
arm design. The 4 degrees of rotation that the arms has enables it to move in any direction, so the
payload can be placed anywhere within the certain distance. The plastic clips have been tested, so
they can secure the payload without letting it move. The design for the REM is simple, but testing
for this part has not begun. The ending angle of the rail will be hard-coded into the Arduino, so
there is less room for error during this phase. Extensive testing will be needed to ensure the
accuracy of this step. The same can also be said about the MIS because the distance the igniter
moves will also be hard-coded into the Arduino. After the REM is finished, the rail needs to
maintain its position at the final angle. In order to do this, the Arduino must constantly monitor
the angle of the rail, in case the rail starts to fall under its own weight.
5.3. Mission Success Criteria and Functional Requirements
Requirement Design Feature Requirement
Verification
Success Criteria
Capture the payload Robotic arm with IR
sensors will locate
and grip the payload
Visual inspection The payload stays in
the grip of the claw
Move the payload
into the payload bay
located in the Launch
Vehicle
Robotic arm will have
4 DOF controlled by
4 servo motors
Visual inspection The arm moves with
speed and stability
and the payload stays
in the grip of the claw
Secure payload in
payload bay
Plastic clips will snap
around the payload
Visual and audio
inspection (from
snapping sound of
plastic clips)
The payload does not
fall out of the pay
Raise the Launch
Vehicle
A cable and spool
system will pull the
guide rail upwards to
appropriate angle
Visual inspection and
touch sensor feedback
The Launch Vehicle
moves from a
horizontal position to
5 degrees from the
vertical

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Maintain the Launch
Vehicle angle
A ratchet system will
ensure the Launch
Vehicle can only
move upwards
Visual inspection and
audio inspection (from
clanking noise of
ratchet locking)
The Launch Vehicle
does not fall down
Insert the igniter A rack and pinion
system will move the
electronic match into
the motor cavity
Visual inspection The igniter is inserted
1 ft. into the motor
cavity
5.4. Manufacturing Approach
5.4.1. Overall Manufacturing Approach
Mission success will be achieved by building a reliable and stable design. Before any actual
manufacturing, the sub-team leader and the team leader must look over the plans. This allows for
any mistakes to be caught before materials are wasted. In order to best achieve this, the parts used
in the AGSE will be manufactured from computer controlled machines, including: water jet, laser
cutter, and 3-D printer. By using these tools, greater accuracy can be achieved than machining by
hand. Parts that cannot be manufactured in house will be purchased from trusted vendors. With a
focus on precise manufacturing, our CAD models will be followed with as little error as possible.
This decreases the chances of an improperly produced part which could lead to failure in a
subsystem.
Using these machines also allows for many parts to be produced at once. These parts can be used
as replacement parts if after testing the AGSE, some parts start to wear out. This ensures a
sustainable design.
5.4.2. RPDS
Accuracy in movement and mechanical strength are the main keys to ensure mission success of
the RPDS. The robotic arm is designed in a manner that its materials will not fail during operation.
The laser cutted plywood struts are positioned with its face sideways to better support the load in
the vertical direction. The laser cutting operation ensures precision in its dimensions and allows
easy and fast manufacturing. Because the servo mounts embody a relatively complex shape, it is
difficult to manufacture through milling operations. As a result, they are 3D printed in high

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quantities, in case any break during operation. The claw is almost entirely made of 3D printed
parts, also because they have complex shapes. The design of the claw is very different from
ordinary claws found in the market. As a result, the design is subject to constant modification,
which is promoted by the easy manufacturing process of 3D printers.
5.4.3. RES
The stability and structural integrity of the RES frame and pulley/cable system are key to mission
success. The frame is both wider and longer than the Launch Vehicle and other necessary
mechanics to increase stability. It is made out of thick T-slotted extrusion beams that are fastened
together using T-nuts. The pulley and stepper motor casing are also attached to the frame’s T-
slotted beams with T-nuts. The pulley is attached to a vertical part of the frame, and the stepper
motor casing is attached to the base of the frame. The stepper motor uses a force fitted coupler to
hold the spool. The cable travels from the spool through the pulley, and is connected to an eye
screw on the lifting rod on the other side of the frame. The connection of the cable to the lifting
rod is crucial to mission success. The cable will be strung through the eye screw, tied and clamped,
to ensure that it is properly connected and does not move from its required position.
5.4.4. MIS
Mission success for the MIS is highly dependent on accuracy. To maintain a high level of accuracy
for the rack and pinion system, all steel pieces will be produced using a computer controlled water
jet. The water jet allows all the pieces to be a uniform shape. After the water jet, the pieces will be
measured manually to confirm the sizes.

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5.4.5. Electronics
For construction of the final product, new wires will be used, and the breadboard itself will be
tested separately to ensure that there are no broken connections. All the connections to the Arduino
and breadboard will be taped to prevent any disconnect, and all the electric components will be in
a metal electric box to prevent further disturbances. A fully charged battery will also be used during
the competition.
5.5. Remaining Manufacturing and Assembly
Subsystem Part Manufacturing
Method
Estimated Completion
Date
AGSE Support frame Bandsaw 2/4
RPDS Servo motor mounts 3-D printing 1/21
RPDS Struts Laser cutting 1/21
RES Stepper motor mount 3-D printing 1/21
MIS Steel guides Water jetting 1/26
MIS Steel sides Water jetting 1/26
MIS Steel blast plate Water jetting 1/26
Electronics Electronics containment
unit
Laser cutting 2/4

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5.6. AGSE Electronics
5.6.1. AGSE Electronics Overview
The AGSE electronics primarily consists of 2 stepper motors, 5 servo motors, and multiple LEDs
and resistors. The bipolar stepper motor will be used for the MIS, the unipolar stepper motor will
be used for the RES, and the RPDS will be comprised of the 5 servo motors. An Arduino Mega
2560-R3 is used as the microcontroller for all 3 processes. The entire system is powered by a 12V
- 10.5Ah lead acid battery. All the electronic components of the AGSE will be housed in a 2ft. x
1.5ft. x 1ft. metal electrical box.

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5.6.2. Drawings and Schematics
Schematic of the AGSE electronics including, motors, resistors, the Arduino, and LEDs. The two
drivers for the stepper motors need to be connected to a breadboard which is not shown. Other than the
servo motors, LEDs, and the stepper motors, every component should be housed in the electronics box as
shown in the schematic.

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5.6.3. Block Diagrams
When the system is powered on and the start button is pressed, the system will activate the RPDS,
which finds and delivers the payload to the rocket. After this process is finished, the green LED
corresponding to the RPDS will turn on, and the system will activate the REM. After the rocket is
raised, the corresponding green LED will turn on, and the MIS will be activated. To ensure the
safety of the AGSE is not compromised, there will be emergency stop switches located around
each subsystem for the AGSE. For the REM, the emergency-stop switch will be located on the
support rail for the main pivot. Figure below displays the emergency stop button location.
A similar system will be used for the MIS whereas once the roller switch underneath the rack
registers a false, the motor will be stopped. The false indicates that the rack is no longer pressing
down on the switch and has moved to its required position.

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5.6.4. Batteries/Power
Component Quantity Total Operating Current
Servo Motor 5 500 mA
Unipolar Stepper Motor 1 2000 mA
Bipolar Stepper Motor 1 330 mA
LED 7 280 mA
Pushbutton 1 40 mA
Arduino Mega 1 100 mA
Based on the previous years’ time: the RES will take 4 minutes, the MIS will take 1 minutes, and
the RPDS will take 3 minutes. Using these times and the current shown above, the amount of
consumed charge can be calculated per full system run-through:
.5(3) + 2(4) + .33(1) + .28(8) + .04(8) + .1(8) = 13.19 / 60 minutes = .2198 Ah for one full run.
The battery contains 10.5 Ah, therefore the battery can power 47.76 runs. This translates to 6.37
hours.
5.6.5. Switch and indicator Wattage and Location
One of the challenges in coding the Arduino was finding a method of pausing and resuming the
code based on the push of a button. The Pause-Button Test was the process used to find the best
method of pausing and resuming the Arduino Code.
The electronics will be housed in a 2ft. x 1.5ft. x 1 metal electronics box. This box will contain the
Arduino, the battery, the breadboard, the resistors, and all the wires. The AGSE Schematic shows
the location of all the components inside the electronics box.

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5.7. Component Testing
5.7.1. Electronics Pause Button Test
The purpose of the Pause-Button Test was to see if a start/pause pushbutton could be implemented
to instantaneously stop the Arduino and continue on another press of the button. The challenge
was getting the Arduino to continue from where it left. off rather than repeat the operation it was
on. This was solved by converting single lines of code into loops. For example, instead of rotating
a servo motor to 100 degrees, the code would incrementally loop to 100 in increments of 1 degree.
Loops allow the Arduino to constantly monitor the state of the button, thereby allowing it to
continue from where it left. off.
After this change was made, the button worked as it should. In the figure above, when the button
was pressed, the servo motor was moving to a certain degree. The servo stopped immediately, and
on the next press, it resumed moving to that same degree. Please refer to Appendix XX, for the
test plan with more details concerning the set up and experimental

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5.7.2. RES Lifting Test
The main purpose of this test is [Appendix XX] to ensure the RES mechanism has the ability to
lift. the launch vehicle, to ensure the cable and framing is safely stabilized, to ensure the individual
components (this includes the stepper motor, the motor casing, the spool, the pulley, the cable, and
the eye screw) are durable enough to withstand the forces involved with lifting the launch vehicle,
to ensure the launch vehicle could be lifted in a safe but efficient rate of speed, and to determine
the most effective position for the pulley to be placed on the frame to yield the most torque on the
lifting rod.
During the testing, the RES will attempt to lift. the launch vehicle as quickly as it can as well as
hold the launch vehicle at different positions (15, 30, 45, 60, 85 degrees) in between the horizontal
and 5 degrees off of the vertical. It will help determine if the cable or frame needs more stability
support. If so, a guide will be added to the cable’s path and more framing will be added to the base
of the existing frame. The height of the pulley position will be varied during testing to determine
optimal torque. If the RES is not able to lift. the launch vehicle because of strength and durability
issues, more pulleys will be added to the frame to help distribute the force of tension. The cable
would then be intertwined between the pulleys, thus making it easier for the RES to lift. the launch
vehicle.
5.7.3. RPDS Payload Insertion Test
The payload insertion test proves the strength capability of the robotic arm and its precision to
accomplish the task. The test is simple and direct. A payload is placed approximately a foot away
from the robotic arm. The arm then approaches the payload, grabs it with the claw, moves it above
the locking component and finally presses it into the lock. This test can be formed in two different
ways. Either through predetermined movements by the servo motors, or through constant feedback
from ultrasonic sensor readings. If the option of the usage of sensors is optimized through trials
and better programming algorithms, it shall be the adequate method for the robotic arm movement.
The ultrasonic sensor will only be used to locate the payload and the locking component.

Page 74 of 92
Because the servo motors’ movements are relatively accurate, it is possible that a hardcoded
movement can outperforms the movement with distance feedbacks. However, the aim is to perfect
the algorithm to search for the payload and its securing component. Additionally, it is expected
that the servo motors used are capable of moving the payload with relative ease. If the servo motors
don’t have enough torque to perform its task, stronger motors are to be used.
5.8. Integration Plan
The AGSE and the Launch Vehicle interface at two specific points, the payload hatch and the
Launch Vehicle rail. The payload hatch interacts directly with the RPDS. The RPDS must be able
to navigate within the hatch to place the payload. To allow this, the hatch will have an opening 2.9
inches wide while the arm has a width of 2.5 in. The RPDS must also have enough strength to snap
the payload into the clips inside the bay. This will be done by adjusting the strength of the clips
after testing. Finally the RPDS must be able to close the hatch. The closing method was simplified
by adding a magnetic lock to the hatch. The RPDS must simply move the hatch forward with its
wrist motion to begin closing and the magnets will finish the closing process.
The other part of the AGSE-Launch Vehicle interface is the launch rail. The Launch
Vehicle will first be secured using rail buttons on the rocket. These rail buttons will prevent the
Launch vehicle from moving perpendicular to the rail. To prevent rotational motion, steel
extrusions along the rail will hold the Launch Vehicle’s fins at a fixed angle. To prevent motion
up and down the rail, a steel plate will be placed near the end of the rail (but in front of the MIS)
to stop the Launch Vehicle from sliding. These additions to the rail will only allow the Launch
Vehicle to move up the rail, restricting unwanted motion.

Page 75 of 92
5.9. Failure and Safety Analysis
Potential Failure Effects of Failure Failure Prevention
Payload is not
secured in bay
Payload will bounce inside
payload bay, disrupting flight
Test various plastic clip dimensions
to find best fit
RPDS stuck inside
payload bay
Payload bay will not close
and RPDS will be destroyed
by raising of the launch
vehicle
RES will be started by a signal from
the RPDS after it has completed its
task
Launch Vehicle
moves uncontrollably
on the rail
Could disrupt performance of
other subsystems
More support along the launch rail to
keep the disruptive movement of the
launch vehicle at a minimum
RES is not stable
while raising
Rocket will not be raised, and
potentially the motors will be
broken
Test subsystem, add counterweights
to reduce necessary force from motor,
and add more framing to increase
stability
RES is not stable at
full extension
Launch vehicle could tip over Increase the weight to lower the
center of gravity. Increase the base
width. Add more supports to the
launch rail.
RES does not stay
upright
Launch vehicle will fall
unpredictably
Perfect ratchet system, ensure tension
in steel cable
RES stepper motor
does not stop
Tension will continue to
increase in the cable leading
to failure
Emergency stop button in place that
activates when rail is at maximum
angle
MIS stepper motor
does not stop
Rack will move further into
motor cavity, possibly
damaging motor
Emergency roller switch in place that
activates when rack passes a certain
distance

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Electronics short
circuit or are
overloaded
System will lose control Fuses will protect electronics,
subscale testing will prevent short
circuits and overloads

Page 77 of 92
6. Project Plan
6.1. Budget Plan
The projected project budget is approximately $5.872.38 – below the projected fundraising goal
by just over 11%. This cost was derived using the actual project costs from the 2015-2016 NASA
SLI competition cycle and a 15% margin was added to the Launch Vehicle and Flight Systems
costs during the previous project cycle. The project budget breakdown is listed numerically in
Table 6: Budget Summary and graphically in Figure 7.
Table 6: Budget Summary
Section Cost
Avionics $700.00
AGSE $808.60
Launch Vehicle$963.78
Testing $900.00
Motor $1,000.00
Operations $1,000.00
Outreach $500.00
Total Budget $5,872.38
Table XX lists the expenses as of the PDR Milestone. The summary is broken down into five (5)
main categories: Launch Vehicle, AGSE, Flight Systems, Operations, and Testing. The Launch
Vehicle and Flight Systems categories are further broken down into two (2) subcategories: Flight
Hardware and Testing. Operational expenses include: non-system specific test equipment, Team
supplies, non-system specific fabrication supplies, as well as any travel and outreach expenses.
Any system-specific equipment bought for testing is charged against that specific system.
$700.00
$808.60
$963.78
$900.00
$1,000.00
$1,000.00
$500.00
2015-2016 ARES Projected Budget Distribution
Avionics
AGSE
Launch Vehicle
Testing
Motor
Operations
Outreach
Figure 7: Budget Summary

Page 78 of 92
Table XX: Expenses as of PDR
Subsystem Amount
Launch Vehicle $75.36
Flight Systems $14.53
AGSE $92.69
Operations $20.00
Testing $2.50
Total $205.08
The Subscale Purchases were roughly around $75.36, include the motor and building the structure.
Many parts for the full scale such as the fiber glass body tube have been purchased but until they
arrive, they will remain unmarked.
6.2. Funding Plan
In order to fund the 2015-2016 competition cycle, Team ARES have sought sponsorships from
academic and industry sources. The current sponsors of Team ARES and their predicted
contributions can be found in Table XX. Additionally, the Team has also received a dedicated
room in which the Team can construct and store their launch vehicle, payload, and other non-
explosive components. All explosive components (i.e. black power) are properly stored in Fire
Lockers in either the Ben T. Zinn Combustion Laboratory or the Ramblin’ Rocket Club Flammable
Safety Cabinet. Furthermore, the Georgia Tech Invention Studio and AE Maker Space will support
all fabrication needs of the Team.
Sponsor Contribution Date
2014-2015 Unused Funds $1,200 --
Georgia Space Grant Consortium $1,000 Nov 2015
Georgia Tech School of Aerospace
Engineering
(est.) $1,000 Jan 2015
Georgia Tech Student Gov’t Association (est.) $1,000 Jan 2015
Corporate Donations (est.) $2,000 Feb 2016

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Orbital ATK Travel Stipend (est.) $400 Apr 2016
Total $6,600

Page 80 of 92
7. Project Schedule
Team ARES project is driven by the design milestones set forth by the NASA SLI Program Office.
The design milestones are listed in Table XX. The project Gantt chart for Project Hermes –
contains only high-level activities due to the unique launch vehicle and payload designs. A more
detailed Critical Path chart is located in Section XX.
Deadline Date
Team Formation 20 AUG
Proposal 11 SEPT
Web Presence Established 23 OCT
PDR Documentation 6 NOV
PDR Teleconference 9-20 NOV
CDR Documentation 15 JAN
CDR Teleconference 19-29 JAN
AGSE, Flight Systems, and
Launch Vehicle Testing
29 JAN –
20 FEB
Full Scale Testing and
Launching
20 FEB
FRR Documentation 14 MAR
FRR Teleconference 17-30 MAR
Competition 13-16 APR
PLAR Documentation 29 APR
7.1.1. Critical Path Chart: CDR to PLAR
The critical path chart illustrated below demonstrates the highly integrated nature of Project
Hermes. The critical path chart identifies:
• High Risk Tasks – red boxes
• Low-Moderate Risk Tasks – pink boxes

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• Earned Value Management (EVM) Goal Tasks – gold boxes
• Nominal Tasks – grey boxes
• Critical Path – green arrow
• Non-Critical Path – black arrows
• Current Place on the Critical path – blue outline

Page 82 of 92
Figure: Critical Path Chart from CDR to PLAR
A larger version of the critical path discussion can be found in the Appendix XX.

Page 83 of 92
7.2. Schedule Risk
Three (3) items have been identified as “High Risk Items.” These are:
• Launch Vehicle Recovery System Design
• Flight Software Verification
• Launch Vehicle & AGSE Systems Design
Table XX lists the mitigation for these items:
CAPTION FOR TABLE XX
High-Risk Task Potential Impact on Project Hermes Mitigation
Launch Vehicle &
AGSE Design,
Fabrication, &
Testing
1) Schedule Impact
2) Budgetary Impact
3) Not qualifying for Competition
Launch
1) Ensure personnel have direct
and free access to
experienced personnel on
and off of the team
2) Ensure personnel have
knowledge to effectively
utilize simulation and
analysis tools
3) Ensure personnel have direct
and free access to the
simulation and analysis tools
4) Ensure personnel are familiar
with relevant fabrication
Recovery System
Design,
Fabrication, &
Testing
1) Excessive kinetic energy during
landing resulting in damage to the
launch vehicle
2) Failure to deploy drogue and/or
main parachute resulting in a high
energy impact with the ground
destroying the Launch Vehicle
1) Ensure Recovery System
Lead has direct and free
access to experienced
personnel on and off the
team
2) Provide real-time feedback
of the design decisions to
ensure all recovery-related
requirements are meet with

Page 84 of 92
at least 5% margin wherever
possible
3) Ensure proper manufacturing
and packing techniques are
utilized during the assembly
and fabrication of the
recovery system
Flight Software
Design,
Verification, &
Testing
1) Incorrect Calculations during
flight leads to an inaccurate
targeting
2) Flight Systems does not function
properly during flight
3) Flight Systems encounters a
flight anomaly that results in
excessive draw and damage to
the Flight Avionics, Power
Supply, and/or Launch Vehicle
1) Ensure Flight Software Lead
has direct and free access to
experienced personnel on
and off the team.
2) Thoroughly test the system
on the ground with wind
tunnels and other testing
3) Develop multiple paths to
achieve end goal of
developing the robust control
logic that is required for the
successful demonstration of
the Flight System

Critical Design Review Paper for Project A.R.E.S.

Critical Design Review Paper for Project A.R.E.S.

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Critical Design Review Paper for Project A.R.E.S.