NASA PDR Technical Report

Tarleton Aeronautical Team
2012 - 2013 USLI Preliminary Design Review
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Note to reader:
To facilitate the reading of the Preliminary Design review, we have mirrored the Student
Launch Project Statement of Work. In the body of the PDR, you will find extensive detail
in the design of our SMD payload. The payload’s features are threefold with
atmospheric data gathering sensors, a self-leveling camera system, and video camera.
One of the two major strengths of our payload design is the originality of our
autonomous real-time camera orientation system (ARTCOS). The other major strength
can be found in the originality of our self-designed Printed Circuit Board layouts. This
feature alone represents over 100 man hours of work. Along with space and power
efficiencies, the PCB’s provide major enhancement of the signal integrity of the sensor
data. For ease of reading, you will find documents such as itemized budgets, and
launch procedures moved to the appendix along with Sensor and Material Safety Data
sheets. We have enjoyed the challenges presented in the writing of this document and
submit it for your review.

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Table of Contents
Table of Contents
List of Figures..................................................................................................................vi
List of Tables...................................................................................................................ix
List of Acronyms..............................................................................................................xi
I) Summary of PDR Report ........................................................................................... 13
Team Summary ......................................................................................................... 13
Launch Vehicle Summary .......................................................................................... 13
Payload Summary...................................................................................................... 13
Launch Vehicle Overview .......................................................................................... 14
Motor Selection....................................................................................................... 14
Recovery ................................................................................................................ 14
II) Changes Made Since Proposal................................................................................. 16
Vehicle Change log................................................................................................. 16
Payload Change Log .............................................................................................. 16
Project Plan Change Log........................................................................................ 17
III) Vehicle Criteria......................................................................................................... 18
Selection, Design, and Verification of Launch Vehicle ............................................... 18
Launch Vehicle Mission Statement......................................................................... 18
Mission Success Criteria ........................................................................................ 18
Propulsion Motor Selection .................................................................................... 26
Performance Characteristics and Verification Metrics................................................ 31
Recovery System.................................................................................................... 31
Structure System .................................................................................................... 32
Propulsion............................................................................................................... 32
Verification Plan...................................................................................................... 32
Risks and Plans for Reducing Risks .......................................................................... 39
Planning of Manufacturing ...................................................................................... 43
Confidence and Maturity of Design............................................................................ 45
Electrical Schematics of the Recovery System.......................................................... 47

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Mass Statement......................................................................................................... 48
Recovery System....................................................................................................... 50
Recovery Component Itemization........................................................................... 61
Mission Performance Predic ...................................................................................... 66
Mission Performance Criteria.................................................................................. 66
Simulations ............................................................................................................. 66
Stability................................................................................................................... 71
Kinetic Energy......................................................................................................... 72
Interfaces and Integration .......................................................................................... 78
Internal Vehicle Interfaces ...................................................................................... 78
Vehicle to Ground Launch System Interfaces......................................................... 81
Launch Operation Procedures ................................................................................... 81
Safety and Environment (Vehicle).............................................................................. 81
The Safety Officer................................................................................................... 81
Failure Modes......................................................................................................... 82
Rocket Design Failure Modes................................................................................. 82
Payload Integration Failure Modes ......................................................................... 83
Launch Operations Failure Modes.......................................................................... 83
Hazard Analysis......................................................................................................... 85
Environment............................................................................................................... 88
Environmental effects of the project........................................................................ 88
Environmental effect on the project ........................................................................ 89
IV) Payload Criteria ....................................................................................................... 90
System Level Review ............................................................................................. 90
Required Subsystems................................................................................................ 94
Atmospheric Data Gathering .................................................................................. 94
Global Positioning System.................................................................................... 104
Wireless Transmitter............................................................................................. 105
Autonomous Camera Orientation System ............................................................ 106
Video Capture....................................................................................................... 111
Liquid Crystal Display ........................................................................................... 112

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Official Scoring Altimeter ...................................................................................... 113
Power Supply........................................................................................................ 114
Performance Characteristics.................................................................................... 114
Verification Plan....................................................................................................... 115
Preliminary Integration Plan..................................................................................... 117
Precision of instrumentation, repeatability of measurement, and recovery system.. 118
Drawings and Electrical Schematics........................................................................ 120
Electrical Schematic................................................................................................. 121
Cross-Component Compatibility .............................................................................. 131
Payload Concept Features and Definition................................................................ 135
Uniqueness and Significance ............................................................................... 137
Suitable Level of Challenge .................................................................................. 137
Science Value.......................................................................................................... 138
Experimental Logic, Approach, and Method of Investigation ................................ 139
Relevance of Expected Data and Accuracy/Error Analysis .................................. 140
V) Project Plan ............................................................................................................ 140
Testing Budget...................................................................................................... 141
Outreach Budget................................................................................................... 141
Travel Budget ....................................................................................................... 142
Budget Summary.................................................................................................. 142
Funding Plan............................................................................................................ 143
Timeline ................................................................................................................... 143
Gantt Timeline ...................................................................................................... 143
Testing Gantt Timeline............................................................................................. 145
Outreach Gantt Timeline....................................................................................... 147
Outreach Plan ............................................................................................................. 148
Educational Outreach .............................................................................................. 148
Educator Outreach................................................................................................... 149
Community Outreach ............................................................................................... 149
Star Party.............................................................................................................. 149
Tarleton Regional Science Olympiad.................................................................... 149

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Participation Goal..................................................................................................... 150
Accomplished Educational Outreach.................................................................... 150
Conclusion .................................................................................................................. 153
Appendix A - Itemized Subsystem Budget .................................................................. 155

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List of Figures
Figure 1 Vehicle Structure............................................................................................. 19
Figure 2 - Nose Cone.................................................................................................... 20
Figure 3 Ballast System ................................................................................................ 21
Figure 4 Upper Body Airframe....................................................................................... 22
Figure 5 Payload Housing............................................................................................. 24
Figure 6 - Fin Design Selection ..................................................................................... 25
Figure 7 - Fin Design Options ....................................................................................... 25
Figure 8 Selected Fin Design ........................................................................................ 26
Figure 9 – Cesaroni L1720 Thrust Curve ...................................................................... 28
Figure 10 - AeroTech L1390 Thrust Curve.................................................................... 29
Figure 11 - Cesaroni L1090 Thrust Curve..................................................................... 30
Figure 12 - Risk Plot...................................................................................................... 42
Figure 13 - Team Hierarchy .......................................................................................... 45
Figure 14 - Project Life Cycle ........................................................................................ 46
Figure 15 - Dimensional Drawings ................................................................................ 46
Figure 16 - Recovery Electrical Schematic.................................................................... 47
Figure 17 - Featherweight Raven 3 Wiring.................................................................... 48
Figure 18 - Flight Sequence.......................................................................................... 50
Figure 19 Daveyfire Electric Match................................................................................ 52
Figure 20 Altitude Simulation ........................................................................................ 67
Figure 21 Velocity Simulation........................................................................................ 68
Figure 22 Acceleration Before Burn Out Simulation...................................................... 69
Figure 23 Acceleration After Burn Out........................................................................... 70
Figure 24 L1720 Thrust Curve ...................................................................................... 71
Figure 25 Stability: Center of Pressure/Gravity ............................................................. 71
Figure 26 Weather Cocking........................................................................................... 76
Figure 27 Weather Cocking Simulation......................................................................... 77
Figure 28 Payload Pre-Integration ................................................................................ 78
Figure 29 Coupler Specifications .................................................................................. 79
Figure 30 Payload Design Configuration....................................................................... 91
Figure 31 Self-leveling Camera Configuration............................................................... 93
Figure 32 Arduino Mega 2560-R3 Microcontroller......................................................... 95
Figure 33 Adafruit Micro SD Adapter............................................................................. 97
Figure 34 BMP180 Breakout Board............................................................................... 98
Figure 35 MS5611-01BA03 Breakout Board................................................................. 98
Figure 36 HIH4030 Breakout Board ............................................................................ 100
Figure 37 HH10D Breakout Board .............................................................................. 100
Figure 38 SP-110 ........................................................................................................ 101
Figure 39 TAOS TSL2561 Breakout Board................................................................. 102

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Figure 40 SU-100........................................................................................................ 103
Figure 41 TOCON_ABC3............................................................................................ 104
Figure 42 LS20031...................................................................................................... 105
Figure 43 XBee-PRO XSC S3B .................................................................................. 106
Figure 44 Arduino Pro Mini 328 Mircocontroller .......................................................... 107
Figure 45 VC0706 Photographic Camera ................................................................... 108
Figure 46 Servo Dimensional Layout .......................................................................... 109
Figure 47 ADXL345 Breakout Board........................................................................... 110
Figure 48 VCC-003-MUVI-BLK Video Camera ........................................................... 111
Figure 49 Sparkfun LCD-11062 Screen ...................................................................... 112
Figure 50 Adept A1E................................................................................................... 113
Figure 51 Atmospheric Data Gathering Conceptual Wiring......................................... 120
Figure 52 Camera Orientation Wiring.......................................................................... 121
Figure 53 PCB Schematic for Payload Sensors.......................................................... 122
Figure 54 I^2C Sensors............................................................................................... 123
Figure 55 Radio Communications ............................................................................... 124
Figure 56 MicroSD Card Reader................................................................................. 125
Figure 57 Analog Sensors........................................................................................... 126
Figure 58 USB Interface.............................................................................................. 127
Figure 59 GPS Module................................................................................................ 128
Figure 60 Voltage Regulators...................................................................................... 129
Figure 61 Preliminary PCB Layout .............................................................................. 130
Figure 62 Component Listing ...................................................................................... 131
Figure 63 SMT (left) vs. Through Hole Devices (right) ................................................ 131
Figure 64 Atmospheric Data Gathering Subsystem Data Flow ................................... 132
Figure 65 Autonomous Camera Orientation System Data Flow.................................. 133
Figure 66 Atmospheric Data Gathering Software Flow Chart...................................... 134
Figure 67 Autonomous Camera Orientation Software Flow Chart .............................. 135
Figure 68 - PCB Board................................................................................................ 136
Figure 69 - Self-Leveling Camera System................................................................... 137
Figure 70 - Initial Funding............................................................................................ 143
Figure 71 Project Timeline .......................................................................................... 144
Figure 72 Testing Timeline.......................................................................................... 146
Figure 73 Outreach Timeline....................................................................................... 147
Figure 74 - Acton Middle School ................................................................................. 148
Figure 75 Subject Interest ........................................................................................... 151
Figure 76 Presentation Learning Outcomes................................................................ 152
Figure 77 Favorite Part................................................................................................ 153

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List of Tables
Table 1 Launch Vehicle Summary ................................................................................ 13
Table 2 Payload Summary............................................................................................ 13
Table 3 Vehicle and Recovery Changes ....................................................................... 16
Table 4 Payload Change............................................................................................... 17
Table 5 Project Plan Changes....................................................................................... 17
Table 6 Upper Body Airframe Trade and Selection....................................................... 21
Table 7 Pros and Cons of Payload Options .................................................................. 23
Table 8 Motor Trade and Selection ............................................................................... 27
Table 9 Vehicle Verification Table................................................................................. 37
Table 10 Recovery System Verification Table............................................................... 39
Table 11 System Risks.................................................................................................. 41
Table 12 Project Risks .................................................................................................. 42
Table 13 Testing Summary ........................................................................................... 43
Table 14 Total Mass Summary ..................................................................................... 48
Table 15 Payload Mass Summary ................................................................................ 49
Table 16 Recovery Mass Summary .............................................................................. 49
Table 17 Structure Mass Summary............................................................................... 50
Table 18 Deployment Altimeter Trade and Selection .................................................... 51
Table 19 Electric Match Trade & Selection ................................................................... 53
Table 20 Recovery Testing Dates ................................................................................. 56
Table 21 Anemometer Trade & Selection ..................................................................... 56
Table 22 Parachute Diameters...................................................................................... 61
Table 23 Main Parachute Trade and Selection ............................................................. 62
Table 24 Drogue Parachute Trade and Selection ......................................................... 62
Table 25 Parachute Protection Materials ...................................................................... 63
Table 26 Shock Chord Trade and Selection.................................................................. 64
Table 27 Launch Day Recovery System Budget........................................................... 66
Table 28 Kinetic Energy Summarization ....................................................................... 74
Table 29 Landing Radius .............................................................................................. 77
Table 30 Ground- Vehicle Interface .............................................................................. 81
Table 31 Potential Failure Modes for the Design of the Vehicle.................................... 83
Table 32 Potential Failure Modes during Payload Integration....................................... 83
Table 33 Potential Failure Modes during Launch .......................................................... 85
Table 34 Potential Hazards to Personnel...................................................................... 87
Table 35 Summary of Legal Risks ................................................................................ 88
Table 36 Effects of Materials used in Construction and Launch.................................... 89
Table 37 Environmental Factors ................................................................................... 90
Table 38 Top Four Candidates for Microcontroller Selection ........................................ 96
Table 39 Top Four Choices for Data Storage Medium.................................................. 96

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Table 40 Micro SD Adapter Selection ........................................................................... 97
Table 41 Pressure Sensor Trade and Selection............................................................ 99
Table 42 Top Considerations for Temperature Sensor Selection.................................. 99
Table 43 Humidity Sensors Trade and Selection ........................................................ 101
Table 44 Pyranometers Trade and Selection .............................................................. 102
Table 45 UV Sensor Trade and Selection................................................................... 104
Table 46 GPS Modules Trade and Selection .............................................................. 105
Table 47 XBee Module Trade and Selection............................................................... 106
Table 48 Microcontrollers Trade and Selection ........................................................... 108
Table 49 Camera Modules Trade and Selection ......................................................... 109
Table 50 Servo Motors Trade and Selection............................................................... 110
Table 51 Accelerometers Trade and Selection ........................................................... 111
Table 52 Top Two Video Cameras Trade and Selection............................................. 112
Table 53 LCD Screen Trade and Selection................................................................. 113
Table 54 Official Scoring Altimeters Trade and Selection ........................................... 114
Table 55 Batteries Trade and Selection ...................................................................... 114
Table 56 Payload Subsystems Evaluation and Verification Metrics ............................ 115
Table 57 SOW Verification.......................................................................................... 117
Table 58 Payload Sensors Precision........................................................................... 119
Table 59 Payload Objectives Summary ...................................................................... 139
Table 60 Proposed Rocket Vehicle Budget Summary................................................. 141
Table 61 Preliminary Testing Budget Summary .......................................................... 141
Table 62 Outreach Budget .......................................................................................... 142
Table 63 Estimated Travel Budget .............................................................................. 142
Table 64 Preliminary Budget Summary....................................................................... 143
Table 65 Accomplished Educational Outreach............................................................ 150
Table 66 Presentation Learning Outcomes................................................................. 151
Table 67 Favorite Part................................................................................................. 152
Table 68 Preliminary Structure Budget............................ Error! Bookmark not defined.
Table 69 Recovery System Budget................................. Error! Bookmark not defined.
Table 70 Payload Budget................................................ Error! Bookmark not defined.

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List of Acronyms
ADC-Analog to Digital Converter
AGL-Above Ground Level
APCP-Ammonium Perchlorate Composite Propellant
APR-Automatic Packet Reporting System
BMP-Barometric Pressure
BOD-Board of Directors
CAD-Computer Aided Drafting
CFR-Critical Design Review
CNC-Computer Numerical Control
DC-Direct Current
EEPROM-Electrically Erasable Programmable Read-Only Memory
EMF-Electric and Magnetic Fields
FAA-Federal Aviation Administration
FRR-Flight Readiness Review
GGA-Global Positioning System Fix Data
GLL-Graphic Position-Latitude/Longitude
GNSS-Global Navigation Satellite System
GPS-Global Positioning System
GSA-GNSS and Active Satellites
GSV-GNSS Satellites in View
HCI-Harris Composites Incorporated
LCD-Liquid Crystal Display
LCS-The Launch Control System
LED-Light Emitting Diode
LSO-Launch Safety Officer
MHz-Mega Hertz
MSDS-Material Safety Data Sheet
N/A-Not Available
NAR-National Association of Rocketry
NASA-National Aeronautics and Space Administration
NMEA-National Marine Electronics Association
NMEA-National Marine Electronics Association
NOAA-National Oceanic and Atmospheric Administration
OSHA-Occupation Safety and Health Administration
PCB-Perforated circuit board
RF-Radio Frequency
RMC-Recommended Minimum Specific GNSS Data
RSO-Range Safety Officer
SD-Secure Digital
SLP-Student Launch Projects
SMD-Science Mission Directorate
SOW-Statement of Work
SPI-Serial Peripheral Interface
STEM-Science Mathematics Engineering Technology

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TAP-Technical Advisor Panel
TRA-Tripoli Rocketry Association
TTL Transistor-transistor logic
USLI-University Student Launch Initiative
UV-Ultraviolet

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I) Summary of PDR Report
Team Summary
Tarleton State University
Box T-0470
Stephenville, Texas 76402
Team Mentor
Our team mentor is Pat Gordzelik.
TRA 5746, L3. TAP NAR 70807 L3CC Committee Chair.
Founder/past Prefect/President Potrocs (Panhandle of Texas Rocketry Society)
Inc., Tripoli Amarillo # 92. BOD member, Tripoli Rocketry Association Inc.
Married to Lauretta Gordzelik, TRA 7217, L2.
Launch Vehicle Summary
Size and Mass
Length 108 inches
Outer Diameter 5.525 inches
Mass 33.5 pounds
Motor
Selection Cesaroni L1720-WT-P
Recovery
Drogue 24” Nylon Parachute, Apogee Deployment
Main 120” Nylon Parachute, 500 foot AGL Deployment
Avionics
Primary Featherweight Raven 3 Altimeter,
Backup PerfectFlite Stratologger Altimeter, and Beeline GPS Tracking
Table 1 Launch Vehicle Summary
The Milestone Review Flysheet can be found in Appendix L.
Payload Summary
Title Experiment
Science Mission
Directorate Payload
Gather Atmospheric and GPS Data, Autonomously Orientate Photographic
Camera, Capture Video for Public Outreach, and a Clear Acrylic Housing
Table 2 Payload Summary

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Launch Vehicle Overview
Size and Mass
The total rocket body is 108 inches in length and has a diameter of 5.5 inches.
There are four body sections: a nose cone, upper body airframe, payload, and a
booster section. The nose cone is fiberglass, consists of an elliptical shape, and
has a 7.5 inch length with a 5.5 inch shoulder. The upper body airframe is 28
inches in length and consists of fiberglass. The payload is 36 inches in length and
is clear acrylic. The booster section is 36 inches in length and consists of
fiberglass. The rocket has a four fin configuration; each fin has a height of five
inches. The simulated unballasted mass of the entire rocket is approximately 33.5
pounds.
Motor Selection
The chosen motor is a Cesaroni L1720-WT-P. The motor has an average thrust of
1754 Newtons and a maximum thrust of 1947 Newtons. The total impulse is 3,696
Newton seconds. It has an initial launch weight of 7.37 pounds and a post burn
weight of 3.5 pounds. This motor is chosen based on simulations taking into
account average launch conditions of the launch site and date. The high initial
thrust is also a factor. The high thrust provides a launch rail exit velocity of 75 feet
per second. With average conditions simulated and no ballast, the predicted
vehicle apogee is 5,342 feet. This apogee, slightly above one mile, allows for the
total mass to increase.
Recovery
The upper body airframe comprises of the main parachute compartment and the
main altimeter compartment. The booster sections include the drogue parachute
compartment and the drogue altimeter compartment. The drogue parachute
deploys at apogee and the main parachute deploys at 500 feet above ground
level. Each deployment altitude is monitored by a redundant system of altimeters,
each is capable of being magnetically locked in the “on” position for the duration of
the flight.
Once apogee is detected, an electronic match charges to ignite a black powder
well which will cause ejection of the drogue parachute. The same mechanism
ejects the main parachute. Both parachutes shield from the heat of the ejection
charges by Nomex parachute bags. Shock harnesses keep the four body sections
tethered after both parachute ejections. Parachute sizes and deployment heights
calculate to ensure that launch vehicle recovery occurs within a 2500 foot radius of

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the launch site. These calculations also ensure that the sections have a maximum
kinetic energy of less than 75 foot pounds force.
Science Mission Directorate Payload
The payload, at a minimum, fulfills the requirements of the SMD payload. This
includes storing and transmitting atmospheric sensor measurements, GPS data,
and taking pictures of the proper orientation. The payload contains the necessary
atmospheric sensors and an autonomous camera orientation system. An onboard
LCD screen assists in pre-launch operations. The payload houses a video camera
to store footage of the flight for educational engagement and public relations
purposes. A clear, acrylic, tube houses the SMD payload. Durable aluminum
mounting rails secure components. The clear acrylic housing allows for internal
solar data gathering, pictures to be taken from within, and visual inspection of
payload components.

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II) Changes Made Since Proposal
In the following sections, tables outline the changes since the proposal. These
reflect outstanding design conflicts that require change in order to meet the
criterion of the project.
Vehicle Change log
As the vehicle design evolves, changes take place and documentation is
performed in order to accommodate healthy communication on the new designs.
The following chart describes the changes thus far in the vehicle design.
Vehicle and Recovery Changes Rationale
Maximum outer diameter from 4 inches to
5.525 inches
Payload section integration and design is better suited for the
wider diameter
Total length from 103 in to 108 in
To allow some additional room for recovery system
implementation
Fin size and design To maintain stability for a larger vehicle
Incorporation of a ballast system in the nose
cone
To allow a greater level of ballast weight control on launch
day
Nose cone length from 7 inches to 7.5 inches Commercial availability after diameter change
Black powder charges more powerful Volume increase within recovery system compartments
Wire mesh added to recovery altimeter
compartments
To provide shielding from wireless transmitting devices
Table 3 Vehicle and Recovery Changes
Payload Change Log
As the current payload design progresses for engineering efficiency, a change log
is kept in order to inform members of changes and facilitate communication. The
following chart recognizes changes pertaining to the payload design.
Payload Changes Reason
Payload width from 3.5” to 5"
To provide more room for autonomous camera
orientation system and inclusion of a video camera
Payload housing will be mounted by screws
through the acrylic to a fiberglass coupler
Avoids manufacturing a threaded acrylic cap

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Inclusion of servos and an accelerometer for the
autonomous camera orientation system
Critical payload objective
PCB planning and design
To reduce overall space required as compared to
perforated board mounting
The official scoring altimeter moved from recovery
bays to payload housing
To avert pressure fluctuations from ejection charges, to
be seen through the clear acrylic
Inclusion of video camera For public relations and educational outreach
Number of humidity sensors from 1 to 2 For redundancy in humidity measurements
Number of pressure sensors from 4 to 2 For efficiency of design
Number of pyranometers from 2 to 4 180 Field of view
Number of UV sensors from 2 to 4 180 Field of view
The top of the payload rail system will rest in a
milled grove of the upper payload mounting cap
For structural support
Two, 90 degree opposing boards for solar
irradiance and UV sensors
Solar irradiance and UV sensors can gather data at 90
degree intervals around the rocket vehicle
Handheld Yagi directional antenna will be used to
receive all transmitted data from the payload
Development of an automated tracking system for the
ground station antenna was deemed unnecessary
Table 4 Payload Change
Project Plan Change Log
As the project develops a change log shall be kept in order to effectively
communicate to the team as a whole the modifications to the project plan. The
chart that follows includes the list of known changes thus far to the project plan.
Project Plan Changes Rationale
Testing Plan and Timeline has been further developed Time Constraints
Educational Outreach Plan has been further developed
To increase quality of Educational outreach
aspect to the entire project
Community Outreach events added
To increase quality of public outreach and
project sustainability
Table 5 Project Plan Changes

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III) Vehicle Criteria
Selection, Design, and Verification of Launch Vehicle
Launch Vehicle Mission Statement
The mission is to design, build, and launch a reusable vehicle capable of delivering
a payload to 5,280 feet above ground level (AGL). The vehicle will carry a
barometric altimeter for official scoring as well as the Science Mission Directorate
(SMD) payload. The design of the vehicle ensures a subsonic flight and must be
recoverable and reusable on the day of the official launch. The launch vehicle
meets the customer prescribed requirements set forth in the Statement of Work
(SOW) of the NASA 2012-2013 Student Launch Projects (SLP) handbook.
Launch Vehicle Requirements
The vehicle adheres to the following primary requirements. The complete list of
requirements is in the Vehicle Verification Table Table 9.
 Vehicle shall carry a scientific or engineering payload. (Requirement 1.1)
 Vehicle shall reach an apogee altitude of 1 mile AGL. (Requirement 1.1)
 Vehicle shall carry one official scoring altimeter. (Requirement 1.2)
 Vehicle must remain subsonic from launch until landing. (Requirement 1.3)
 Vehicle must be recoverable from a 2500 foot radius away from the launch
pad and reusable on the day of the official launch. (Requirement 2.3)
 Vehicle must use a commercially available APCP motor with no more than
5,120 Newton-seconds of impulse. (Requirement 1.11, 1.12)
Mission Success Criteria
The project defines a successful mission as a flight with payload, where the
vehicle and SMD payload are recovered and able to be reused on the day of the
official launch. Moreover, the vehicle will not exceed 5,600 feet and the official
scoring altimeter will be intact and report the official altitude. The recovery system
stages a deployment of the drogue parachute at apogee and follows deployment
of the main parachute at 500 feet. After apogee and descent, the entire vehicle
lands within 2,500 feet of the launch pad.
System Level Review
This section reviews the design of the vehicle which includes structure, propulsion,
and recovery. The project requires the team to consider, research, and analyze

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various concepts. Each system contributes to the overall success of the mission
and helps determine how useful each system is.
Trade and selection is performed for the structure and propulsion systems. The
Recovery Subsystem contains recovery, trade, and selection. Calculations and
measurements for each system are a part of the presentation.
Structure
Selection of vehicle components contains variables of durability, aerodynamics,
cost, availability, and functionality. The structure of the vehicle composes of the
nose cone, upper body/lower body airframe, payload section, tube coupling, and
fin structure.
The structure is able to withstand the substantial forces throughout the flight. It
must also remain subsonic and be reusable. Material, size, and design must be
adequately chosen to fulfill many requirements. The material composition is
chosen primarily to remain reusable. The elliptical nose cone, four fin design, and
material composition of the rocket all contribute to remaining subsonic throughout
the entire flight. The material composition of the payload section is chosen to be
transparent. The fins are also chosen for their aerodynamic properties. The
diameter must be able to house the SMD payload and parachute deployment
systems. The coupling system was designed to couple two different sizes of inside
diameters and to ensure separation or non-separation. Figure 1 displays the
structure of the vehicle.
Figure 1 Vehicle Structure
Nose Cone
The design is an elliptical nose cone. The vehicle is aerodynamic and remains

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subsonic as per the requirements for the flight. The four types of nose cones for
considerations are elliptical, parabolic, conical, and Von Karman. At subsonic
speeds, elliptical and parabolic nose cones generate the least amount of drag. The
conical and Von Karman style nose cones are longer and add excess weight and
more drag at subsonic speeds, so they are not well fit for this project. The
parabolic nose cone experiences slightly less drag but has slightly more weight in
comparison to the elliptical nose cone. An elliptical nose cone is ideal due to the
commercial availability and manufacturing cost of a parabolic nose cone.
Figure 2 - Nose Cone
The nose cone in Figure 2 above is 13 inches long including the 5.5 inch shoulder,
and is 0.075 inch thick. These dimensions accommodate the ballast system, seen
in Figure 3. The ballast system consists of two bulkheads, one higher in the nose
cone that is not removable and one at the exit of the shoulder that is removable. A
five inch bolt stretches between the bulkheads and washers are added to adjust
the ballast mass.

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Figure 3 Ballast System
Upper Body Airframe/Booster Section
Fiberglass is the material of choice for the upper body airframe and booster
section. The durability of fiberglass improves the chances of the rocket being
reusable (Requirement 1.4). Fiberglass is also readily available.
The upper body airframe houses the main parachute deployment system. The
lower body airframe houses the drogue parachute deployment system, the motor
mount, and secures the fins. These components need to withstand all stresses
present during flight. Three materials are considerations for the upper and lower
body airframes: blue tube 2.0, fiberglass, and glass phenolic tubing. Table 6
represents a trade and selection of the component materials in the upper body
airframe.
Material
Peak Load
(lbf)
Peak Stress
(psi)
Modulus
Avg. Cost per
Foot
Availability
(1 highest)
Fiberglass 19256.1 37806.2 2980.8 $38.84 1
Blue Tube 3211.1 5293.4 607.1 $13.74 3
Glassed Phenolic 7758.9 8983.4 1228.9 $50.27 2
Table 6 Upper Body Airframe Trade and Selection

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The upper body airframe is 28 inches long and the lower body airframe is 36
inches long. Both are 0.075 inches thick. These lengths provide adequate space
for the deployment systems. Figure 4 depicts the upper body airframe with nose.
The booster
Figure 4 Upper Body Airframe
Payload Structure
The payload housing structure consists of clear acrylic. This section houses the
SMD payload. Considerations exist for three options for the payload housing
structure. Option one was an independent fiberglass section. Option two consists
of an extension of the upper body airframe. Option three is an independent acrylic
section.
Payload Airframe
Options
Separate Fiberglass
Airframe
Extended Upper
Airframe
Acrylic
Airframe

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The first option is to couple the payload housing structure to the upper airframe.
The payload is removable in order to inspect electronics. The sensors for detecting
ultraviolet radiation and solar irradiance have to be mounted on the exterior of the
vehicle.
The second option is to have one upper body section that contains the main
parachute deployment system in the first half and the SMD payload in the second
half. This reduces the amount of weight because of the lack of a second coupler.
Sensors would have to mount on the exterior and electronics checked through
dismantling.
The third option is an acrylic payload section. This allows all sensors and the
camera to mount internally. Also, this allows us to visually check that all electronics
are functional. Acrylic is also chosen because of its similar properties to fiberglass.
This option does not require the payload to be deployable from the vehicle to
gather data. This is safer and eliminates the chance of failure upon ejection of the
payload.
The material selected for the payload housing structure is cast acrylic. This is the
best option for the electronics inspection, camera and sensor functionality and
safety. The payload section, seen in Figure 5, is 36 inches long, 5.5 inches wide,
and 0.125 inches thick. The increase in thickness of the acrylic is chosen to ensure
adequate strength for the payload section. The compressive properties of acrylic
are undergoing testing with standard ASTM methods to verify the integrity of this
selection. The acrylic tube manufacturer provides a data sheet on this selection
(refer to Appendix K). The compressive yield strength of acrylic, according to
MatWeb (Material Property Data website), is 18,000 psi. The force the rocket
Pros
Durability, cost,
availability, familiarity
Durability, cost,
availability, familiarity,
lack of coupling system
(lighter)
Visibility of electronics
and payload
components from
exterior, internally
mounted sensors
Cons
Inability of external
inspection, externally
mounted sensors,
coupling system
Inability of external
inspection, externally
mounted sensors,
inaccessibility to
internal systems
Durability, weight
Table 7 Pros and Cons of Payload Options

24
motor exerts is 386 pounds and the weight of the upper body airframe and nose
cone is 7.42 pounds. Therefore, the vehicle is able to withstand loading during the
flight. Additional testing at a materials testing facility, as well as full scale pending
launch testing provide more accurate data on the acrylic payload.
Figure 5 Payload Housing
Airframe Coupling
The coupling system must keep vehicle sections together during flight and
separate as needed. The vehicle design requires coupling the upper body airframe
to the acrylic payload structure. These have different inside diameters. The design
provides for a single coupler to fit both. It is 11 inches long, allowing 5.5 inches of
insertion into each section. The outside diameter of the coupler is 5.25 inches to
meet the acrylic inside diameter. The inside diameter of the fiberglass upper
airframe is 5.375 inches. Thus, 5.5 inches of the coupler wrap in fiberglass and
resin to increase the outside diameter to 5.375 inches. Since the upper body
section does not need to separate from the payload section, these sections are
rivet to the coupler. Rivets are superior to epoxy in fastening and allow the area
inside the coupler to be accessible by a user by removing the rivets.
The lower body sections are friction fitted to allow separation with the black
powder charges. This coupler is 11 inches in length, allowing 5.5 inches of
insertion into each section.
Fin Structure
Figure 6 is an exact dimensional drawing of the fin design selection.

25
Figure 6 - Fin Design Selection
The fins are the most important component to flight stability. They determine the
center of pressure, create drag, and create the corrective moment force for
stability. The primary option for fins is three versus four fins. Three fins provide
less drag, less weight, and less corrective moment force. The change in stability
margin between three fins and four fins is insignificant because the weight added
moves back the center of pressure and center of gravity nearly equal amounts.
Although the four fins weigh more, they create more corrective moment force for
stability. The four fins also create more drag on the vehicle, which is desired for
the motor choice.
Using Open Rocket to simulate the vehicle design, the fin design simulations rend
these the findings in Figure 7:
Alternative Option 2Alternative Option 1Selected Option
Figure 7 - Fin Design Options

26
The selected option achieves proper altitude and provides a good stability margin.
The alternative option one performs well, but the reusability of the vehicle could be
compromised. This option could result in structural damage upon landing, as the
fins extend beyond the base of the vehicle. Alternative option two increases drag,
resulting in the vehicle failing to reach altitude with the selected motor. According
to data from simulations, the selected option is chosen in a four fin design as seen
in Figure 8.
Root Chord: 12 inches
Tip Chord: 0 inches
Height: 5 inches
Sweep Length: 9.8 in
Sweep Angle: 63 degrees
Propulsion
Motor Selection
The selected motor is a Cesaroni L1720-WT-P. The high initial thrust helps to
stabilize the rocket as it departs from the launch rail. Its high thrust to weight ratio
is also beneficial to stability. Through simulations that take into consideration the
average conditions for the launch site and date, the Cesaroni L1720-WT-P causes
the vehicle to achieve an apogee of one mile AGL. A ballast system alters the
rocket’s weight to account for alterations between conditions on launch date and
simulated conditions to ensure apogee height of one mile AGL.
The three viable options for motors are Aerotech L1390G, Cesaroni L1090SS-P,
and Cesaroni L1720-WT-P. Each of the motors has a diameter of 2.95 inches.
OpenRocket simulates all motor options. Motors options must achieve apogee
above one mile allowing for the incorporation of ballast weight as well as the
increase to vehicle mass as the design evolves. The Aerotech L1390G motor
results in an apogee height of 5618 feet and off the rail velocity of 65.6 feet per
second, which is acceptable but not optimal. The Cesaroni L1720-WT-P achieved
just over one mile, which allows for weight increase and ballast if needed. Its
Figure 8 Selected Fin Design

27
velocity off the rail is 75.9 feet per second, providing adequate stability. The
Cesaroni L1090SS-P achieves an apogee of 1200 feet above one mile, and a fully
ballasted configuration will not achieve the desired altitude. Table 8 is a trade and
selection table of the motor options.
Motor
Apoge
e (ft.)
Velocity
Off Rail
(ft./s)
Total
Impulse
Max.
Velocity
(ft./s)
Average
Thrust
Burn
Time
(s)
Thrust
to
Weight
Ratio
Availability
/
Cost
Cesaroni
L1720-
WT-P
5345 75.9
830.9lbf
s
(3696Ns
)
738
394.3lbf
(1754N)
2.15 11.8
High/
$170.96
Aerotech
L1390G
5618 65.6
887.8lbf
s
(3949Ns
)
723
308.9lbf
(1374N)
2.65 9.2
Medium/
$209.99
Cesaroni
L1090SS-
P
6479 76.8
1082lbfs
(4815Ns
)
733
246.6lbf
(1097N)
4.4 7.4
Medium/
$346.95
Table 8 Motor Trade and Selection
Cesaroni L1720
The Cesaroni L1720 has a total impulse of 3696 Newton-seconds, which does not
exceed the total impulse maximum of 5120 Newton-seconds. The motor’s
corresponding thrust curve as calculated by Rocksim software is represented in
Figure 9. As shown in the thrust curve, the motor has a fairly neutral motor burn.
As shown in Table 8 and marked in Figure 9, average thrust for this motor is
394.3lbf = 1754N. With this motor, the launch mass of the rocket is 536oz =
15.2kg. Noting that the acceleration of gravity is approximately 9.8m/s², for this
motor the thrust to weight ratio is achievable by = 11.8 : 1, which
exceeds the suggested ratio of 5 : 1.

28
Figure 9 – Cesaroni L1720 Thrust Curve
Aerotech L1390G
The Aerotech L1390G has a total impulse of 3949 Newton-seconds, which does
not exceed the total impulse maximum of 5120 Newton-seconds. The motor’s
corresponding thrust curve as calculated by the Rocksim software is represented
in Figure 10. As shown in the thrust curve, the motor begins with a progressive
motor burn reaches maximum thrust and then begins a regressive motor burn. As
shown in the above table and marked in thrust curve, average thrust for this motor
is 308.9lbf = 1374N. In the following calculation, the mass of the rocket at launch is
used because it represents the maximum mass that the motor would have to be in
order to lift during the flight. In order that the motor be able to lift the rocket, it must
produce enough thrust to overcome the force of gravity, or enough mechanical
energy to achieve a thrust to weight ratio of at least 1.0. In general for a high-
powered rocket, the thrust to weight ratio is given by
( ) ( )
.

29
With this motor, the launch mass of the rocket is 557.3oz = 15.8kg. Noting that the
acceleration of gravity is approximately 9.8m/s², for this motor the thrust to weight
ratio is given by = 9.2 : 1, which exceeds the suggested ratio of 5 :
1.
Figure 10 - AeroTech L1390 Thrust Curve
Cesaroni L1090
The Cesaroni L1090 has a total impulse is 4815 Newton-seconds, which does not
exceed the total impulse maximum of 5120 Newton-seconds. The motor’s
corresponding thrust curve as calculated by the open rocket software is
represented in Figure 11. As shown in the thrust curve, the motor quickly reaches
the maximum thrust then starts a regressive motor burn. As shown in Table 8 on
page 27 and marked in Figure 11, average thrust for this motor is 246.6lbf =
1097N. With this motor, the launch mass of the rocket is 610.2oz = 17.3kg. Noting
that the acceleration of gravity is approximately 9.8m/s², for this motor the thrust to
weight ratio is given by = 6.47 : 1, which exceeds the suggested
ratio of 5 : 1.

30
Figure 11 - Cesaroni L1090 Thrust Curve
Motor Retainer
The motor retainer needs to withstand the massive force of the motor on the
rocket. Three commercially available motor retainers are considerable options.
The first is a quick release system using a cap that snaps onto the retainer body.
The second option is the same retainer body design, only with the implementation
of a threaded cap. Both of these options are simply glued to motor tube. The third
option includes a flange around the retainer body that allows for 12 screws to
mount the retainer to the lower centering ring of the motor tube. It is also glued to
the motor tube and uses a threaded cap for securing the motor in place. This
appears to be the optimal option because of the added protection of mounting to
the centering ring.
Recovery System
The recovery subsystem contributes to the overall mission by ensuring that the
vehicle lands in a reusable condition within a 2500 foot radius from the launch site.
Landing in a completely reusable condition entails that the payload and all other
electronic and mechanical components remain in sound condition throughout the
flight, including impact. To ensure that no component sustains irreparable damage

31
during ascent, the recovery electronics are mechanically armed and programmed
to eject the drogue and main parachutes at safe altitudes.
The ejection charges are sealed off and triggered electronically by redundant
altimeter systems. All bulkheads and materials acted on by the charge firing are
shown through testing to be sufficiently durable to withstand both the energetic
impact and the heat of separation and ejection. Ensuring that the payload remain
functional as a result of impact allows for the continual transmission of data upon
landing, some of which will aid in physically locating the vehicle for a full recovery.
Performance Characteristics and Verification Metrics
Recovery System
The performance of the recovery system relates to the vehicle’s ability to safely
return to the ground. The recovery system must manage the speed of the vehicle
in order to keep the kinetic energy of the each section below 75 ft-lbf. A kinetic
energy greater than 75 ft-lbf could result in the structural failure and cause the
vehicle to be non-reusable. To perform evaluation of parachute size, these
equations are necessary:
∑
These equations are useful during multiple full scale test launches. Performance of
the recovery system also depends on the correct operation of the deployment
altimeters. Necessary evaluation of the recovery altimeters comes through
research and design of electrical wiring diagrams. Verification of performance
comes through ground testing and test launches.

32
Structure System
The performance of the structure system is characterized by the ability to
effectively allow the vehicle to be reusable and to efficiently integrate each
subsystem. In the structure design are the materials in use and fin design, which
contribute to the flight stability of the vehicle. The performance of the fins affects
the entire rocket in that if they fail, the flight becomes unstable and possibly
unsuccessful. Unsuccessful material choice or integration leads to failure if the
rocket cannot withstand the forces of the motor upon launch and the force of
impact with the ground upon landing. Materials’ evaluation is achievable through
material properties databases and the undertaking of tests to withstand the force
of the motor. It withstands testing from the Harris Composite Inc. material testing
facility. The selected materials are then fully verifiable through the full scale test
launches.
Propulsion
The performance characteristic of the propulsion subsystem lies in how
consistently the motor performs so that flight predictions calculate accurately.
Research and simulation achieve the evaluation of the propulsion system. The
motor retainers must withstand the force of the motor, and failure results in entire
mission failure. Both the motor and the retainer undergo static testing to ensure
accuracy and strength. Full scale test launches provide data for verification. Trade
and selection is viewable in Table 8.
Verification Plan
The verification plan in effect reflects how each requirement to the vehicle and
recovery system satisfies its function. Requirements from the SOW are listed and
paraphrased, followed by the satisfying feature of the design to that requirement.
Ultimately, each design feature undergoes verification to ensure that it actually
meets its requirements. Testing, analysis, and inspection serves as the mode of
verification for each feature. A detailed Gantt chart containing test dates is in
Figure 62.

33
Requirement
(SOW)
Vehicle Requirement
Satisfying Design
Feature
Verification
Method
1.1
Vehicle shall deliver payload to
5,280 feet AGL
Motor selection
Testing,
Analysis
1.2
Vehicle shall carry one official
scoring barometric altimeter
Altimeter Model X is used
and included in the SMD
payload section
Inspection
1.2.1
Official scoring altimeter shall
report the official competition
altitude via a series of beeps
Altimeter Model X has
this functionality
Testing,
Inspection
1.2.2
Teams may have additional
altimeters
Four additional
altimeters, outside of the
payload, will be used to
detect apogee and ignite
ejection charges
Inspection
1.2.2.1
At Launch Readiness Review, a
NASA official will mark the
altimeter to be used for scoring
Official altimeter
placement allows ease of
locating and marking
Inspection
1.2.2.2
At launch field, a NASA official
will obtain altitude by listening to
beeps reported by altimeter
The official altimeter has
this functionality
Testing,
Inspection
1.2.2.3
At launch field, all audible
electronics except for scoring
altimeter shall be capable to turn
off
No other electronics in
the design have audible
indicators
Inspection

34
1.2.3.1
Official, marked altimeter is
damaged and/or does not report
an altitude with a series of beeps
Functional Recovery
System
Testing,
Inspection
1.2.3.2
Team does not report to NASA
official designated to record
altitude with official marked
altimeter on launch day
This task will be assigned
to an appropriate team
member
Analysis,
Inspection
1.2.3.3
Altimeter reports apogee altitude
of over 5,600 feet
Motor selection
Testing,
Analysis
1.3
Launch vehicle remains subsonic
from launch until landing
Motor selection
Testing,
Analysis
1.4
Vehicle must be recoverable and
reusable
Recovery system allows
a safe landing of vehicle
Testing,
Inspection,
Analysis
1.5
Launch vehicle shall have a
maximum of four independent
sections
Vehicle is composed of 3
tethered sections
Inspection
1.6
Launch vehicle shall be prepared
for flight at launch site within 2
hours
Launch operations and
assembly procedure
Testing,
Inspection

35
1.7
Launch vehicle will remain
launch-ready for a minimum of
one hour with critical functionality
All critical on-board
components will have
sufficient capacity to
meet entire system
runtime (1.5 hours)
Testing,
Inspection,
Analysis
1.8
Vehicle shall be compatible with
either 8 feet long 1 inch rail
(1010)
1010 rail buttons
attached to vehicle body
Inspection
1.9
Launch vehicle will be launched
by a standard 12 volt DC firing
system
Motor selection is
compatible with this firing
system
Inspection
1.10
Launch vehicle shall require no
external circuitry or special
equipment to initiate launch
Motor ignition only
requires the 12V DC
firing system
Inspection
1.11
Launch vehicle shall use a
commercially available, certified
APCP motor
Cesaroni L1720 Inspection
1.12
Total impulse provided by launch
vehicle will not exceed 5,120
Newton-seconds
3695.6 Ns Inspection
1.15
The full scale rocket, in final flight
configuration, must be
successfully launched and
recovered prior to FRR
Testing Schedule Testing
1.15.1
Vehicle and recovery system
function as intended
Featherweight and
Stratologger Altimeters,
Parachute Calculations
Inspection
1.15.2
Payload does not have to be
flown during full-scale test flight.
The schedule allows for
the payload to be flown in
the full scale launch.
Schedule

36
1.15.2.1
If payload is not flown, mass
simulators shall be used to
simulate payload mass
If this occurs mass will be
added proper sections.
Testing
1.15.2.1.1
Mass simulators shall be located
in same location on rocket as the
missing payload mass
Mass of each section is
calculated, and mass will
be added in the
appropriate sections.
Testing
1.15.2.2
Any energy management system
or external changes to the
surface of the rocket shall be
active in full scale flight
There will be no changes
to the external surface of
the rocket.
Design
1.15.2.3
Unmanned aerial vehicles, and/or
recovery systems that control
flight path of vehicle, will fly as
designed during full scale
demonstration flight
N/A N/A
1.15.3
Full scale motor does not have to
be flown during full scale test
flight
The schedule and budget
plan for the full scale
motor to be flown.
Schedule
1.15.4
Vehicle shall be flown in fully
ballasted configuration during full
scale test flight
The schedule plan for the
fully ballasted system.
Schedule
1.15.5
Success of full scale
demonstration flight shall be
documented on flight certification
form, by a Level 2 or Level 3
NAR/TRA observer, and
documented in FRR package
Pat Gordezlick will be
present at the full scale
launch.
Schedule
1.15.6
After successfully completing full-
scale demonstration flight, launch
vehicle or any components shall
not be modified without
concurrence of the NASA Range
Safety Officer (RSO)
The schedule plans for
the design to be
complete and changes to
be complete.
Schedule
1.16
Maximum amount teams may
spend on rocket and payload is
$5000
Budget indicates that the
total spent on the rocket
and payload is less than
$5000.
Inspection,
Analysis

37
Table 9 Vehicle Verification Table
Table 10 is the verification table for the recovery system.
Requirement
(SOW)
Recovery System
Requirement
Satisfying Design
Feature
Verification
Method
2.1
Recovery devices shall be
staged such that a drogue
parachute is deployed at
apogee; main parachute is at
500ft.
Altimeters will stage
ejection charges for
respective parachutes at
the prescribed altitudes
Testing
2.2
Each independent section of
the launch vehicle will have a
maximum KE of 75 ft-lbf.
Main parachute
selection
Testing, Analysis
2.3
Each independent section of
the vehicle shall land with
2500 ft. of the launch pad
Drogue and main
parachute selection
Testing, Analysis
2.4
Recovery electrical circuits
shall be independent of
payload electronics
Recovery circuits are
independent of payload
with dedicated power
supplies
Inspection
2.5
Recovery system must
include redundant altimeters
Each deployment event
is controlled by a main
and backup altimeter
Inspection
2.6
Each altimeter shall be
armed in launch configuration
with external arming switches
Port holes in vehicle
airframe to altimeter
bays
Inspection

38
2.7
Each altimeter shall have a
dedicated power supply
Each altimeter uses a
separate 9 Volt battery
Inspection
2.8
Each arming switch can be
locked in the ON position
Port holes in vehicle
airframe to altimeter
bays
Testing,
Inspection
2.9
Each arming switch will be a
max of 6ft above the vehicle
base
Main altimeter bay is
located at 6 feet above
the base; drogue
altimeter bay is located
at 2 feet 4.5 inches from
the base
Inspection
2. 10
Removable shear pins shall
be used for the drogue and
main parachutes
compartments
Nylon shear pins will be
used to couple
parachute
compartments
Inspection
2.11
The launch vehicle must
have an electronic tracking
device
2.12
Recovery system electronics
shall not be adversely
affected by any other on-
board electronic device
Altimeter compartments
are shielded by intenal
copper mesh lining
Inspection,
Testing
2.12.1
Altimeters for the recovery
system must be in a separate
compartment than any other
transmitting device
Each altimeter bay has
a dedicated and
separate compartment
Inspection

39
2.12.2
Recovery electronics shall be
shielded from all on-board
transmitting devices
copper mesh lining
Inspection,Testing
2.12.3
shielded from any magnetic
waves generated by on-
board devices
copper mesh lining
Inspection,
Testing
2.12.4
shielded from any on-board
device that could adversely
affect proper operation
copper mesh lining;
Each altimeter bay has
a dedicated and
separate compartment
Inspection,Testing
2.13
Recovery system shall use
commercially available low-
current e matches for ignition
of ejection charges
Davyfire N28BR e-
matches have been
selected
Inspection,
Testing
Table 10 Recovery System Verification Table
Risks and Plans for Reducing Risks
System Risks:
Each system has specialized risks. Certain risks are more likely to occur than
others, and some risks have a more severe consequence. In order to avoid the
realization of a risk, mitigation is performed. Table 11 lists some of the specific
risks to each system, the risk’s likelihood, severity, consequence, and mitigation.
System Risk Likelihood Severity Consequence Mitigation
Safety
Unable to
Obtain Flight
Waivers Medium Medium
Delay in Test
Launches
Schedule Test
Flight at
Official Test
Launch Sites
Disregarding
Safety Plan Low High
Harm to
Participants
Strict Safety
Plan

40
Enforcement
Ignorance of
Legal Rules and
Regulations Low High
Legal
Repercussions
Recurring Legal
Education
Structure
Destruction of
Parts During
Testing High Low
Reordering of
Parts, Delay in
Project
Inventory Extra
Parts
Improper Test
Vehicle
Assembly Medium Medium
Unpredictable
Performance
Assembly
Checklist
Propulsion
Inconsistent
Motor
Construction Low High
Unpredictable
Performance
Thorough
Inspection and
Analysis
Incorrect
Motor Mount
Assembly Low High
Possible Harm
to Vehicle and
Participants
Assembly
Checklist
Educational
Engagement
Poor Quality
Presentation Low High
Less Effective
Outreach
Rehearsed
Presentation
Scheduling
Conflict With
The Schools Medium High
Cancellation of
Events
Communicate
With Schools
Electronics
Faulty
Components Low Medium
Delay to
Project,
Ineffective
Payload
Ordering
Duplicates,
Testing
Incorrect
Wiring
Configurations Low Low
Harm to Parts,
Ineffective
Payload
Thorough
Research and
Design
Power Budget
Miscalculation Low High
Harm to Parts,
Ineffective
Payload
Redundant
Calculations
Recovery
Incorrect Black
Powder Rating Low High
Harm to
Vehicle and
Participants
Extremely
Thorough
Testing
Parachute Size
Miscalculation Low Low
Violation of
Competition
Rules
Redundant
Calculations
Public
Relations
Lack of Project
Exposure Low Medium
Lack of
Funding and
Support
Constant
Outflow of
Updates and
Information
Harmful
Representation
of Team Low Medium
Lack of
Funding and
Support
Proper Ethics
and
Professionalis
m
Management Inadequate High Medium Delay in Weekly Team

41
Communicatio
n
Project, Project
Failure, Design
Flaws
Meetings,
Email Minutes
Scheduling
Conflicts Medium Medium
Delay in
Project
Maintain
Scheduling
University
Administration
Delays Medium Low
Delay in
Project/Fundin
g
Proper
Communicatio
n and
Scheduling
Software
Loss of Source
Code Medium High
Rewrite
Software
Backup Source
Code
Incorrect
Algorithm
Design Low Medium
Ineffective
Payload Logic
Redundant
Calculations
Incorrect
Datasheets Low Low
Confusion,
Delay in
Project
Datasheet
Verification,
Testing
Table 11 System Risks
Project Risks:
There are many risks to the overall project. Each risk has a certain level of
likelihood and severity. The consequences of such risks effect the overall
operation of the team. It is important to mitigate such risks and decrease their
likelihood. Figure 12 displays the risk plot, where all risks in consideration above
plot according to the likelihood they will occur versus the severity of their
consequence. Table 12 lists each risk with an associated number on the plot such
that every risk is identifiable on the plot.

42
Severity # Severity Probabilty Severity # Severity Probabilty Severity # Severity Probability
LOW Medium High
20 Delay in Deliveries 90 60 Communication Failure 10 90 InadequatePersonnel 25
33 Destruction of Parts During Testing 85 66 Unableto Obtain FlightWaivers 0 80 Manufacturing Issues 20
1 IncorrectWiring Configurations 50 45 Improper TestVehicleAssembly 15 99 NAR/TRAViolations 5
3 ParachuteSizeMiscalculation 10 39 Faulty Components 4 98 Damageof Property 15
30 University Administration Delays 40 33 Lack of ProjectExposure 20 70 OSHAViolations 10
5 IncorrectDatasheets 5 34 Harmful Representation of Team 10 100 Personal Injury 1
50 InadequateCommunication 66 67 Teammates Disregarding Safety Plan 2
55 Scheduling Conflicts 75 99
Ignoranceof Legal Rules and
Regulations 3
35 IncorrectAlgorithmDesign 55 72 InconsistentMotor Construction 22
57 EnvironmentPrevents Recovery 25 75 Loss of SourceCode 50
52 Poor Quality Presentation 10 33 Poor Weather 18.5
29 Scheduling ConflictWith TheSchools 12 66 Burn Ban in Effect 25
42 Power BudgetMiscalculation 1 92 Loss of RocketLab 5
93
Loss of Low-AltitudeTestLaunch
Facility in Glen Rose,TX 5
94
Loss of High-AltitudeTestLaunch
Facility in Cross Plains,TX 5
38 Loss of ScienceBuilding 0
39 Engineering Building 0
40 Loss of HCI facilities 12
Figure 12 - Risk Plot
Table 12 Project Risks

43
Planning of Manufacturing
The team has two Industrial Technology majors, experienced in detailed CAD
design including 3D modeling as well as manufacturing/machining with CNC
machines. The team uses the university’s manufacturing facilities to mill, drill, cut,
lathe or machine parts. Members have proper training in the safe operation of such
machines. Access to this facility will allow the team to make adjustments to the
structure design. This is a time efficient alternative to ordering manufactured
structure components.
Planning of Verification
The verification table (Table 9, pages 32-36) indicates all respective requirements
processes for verification. This table guides the group throughout all testing
phases, and all test data analysis ensures that all requirements are in a state of
satisfaction. The Testing Summary Table (Table 13) illustrates a summary of the
testing that is pending. For dates and deadlines of verification testing, refer to the
project plan section with the testing Gantt chart, Figure 60.
Testing Title Subsystem
Structure Testing Structure
Lab Prototyping Structure, Recovery, Integration
Low Altitude flight Structure, Recovery
Dual deployment Recovery
Force of impact Structure, Recovery, Integration
Full scale launch Vehicle, Recovery, Integration
Timed final assembly Structure, Recovery, Integration
Table 13 Testing Summary
Structure testing is at Harris Composites’ testing facilities in Granbury, TX. This
company performs materials strength testing. Arrangements exist to test the
materials that make up the vehicle’s airframe. In particular, they provide data on
material strength and integrity. These figures verify vehicle performance under
expected loads.
Lab prototyping takes place on all subsystems of the vehicle to ensure
compatibility and feasibility, as well as to identify any immediate flaws in the design
or manufacturing. This includes bench top testing, representative model and
prototype builds, as well as documenting and modifying the changes to the design
as needed.
Low altitude flights take place for proof of concept performance, where critical

44
functionality of the vehicle can be verified. Motors for this test are less expensive
than a full scale launch, and flight waivers are more readily available. This allows
adequate test flights to take place and sufficient test data to generate for
diagnostics. The relative ease in conducting this low altitude flights makes this a
very valuable test mode.
Dual deployment testing takes place to ensure functionality of the recovery
system. This includes testing on the ground to verify separation events and
parachute ejection, as well as perfecting ejection charge specifications. Dual
deployment is in the plan for all test flights, both low altitude and full scale.
Force of impact testing will take place to analyze and verify the structural
integrity of the vehicle at landing, as well as the functionality of the recovery
system. The team employs a testing accelerometer in drop testing, as well as in
test flights. This will provide a quantitative justification that the recovery system is
sufficient to meet all requirements and the structural design and integration of the
vehicle is adequate.
Full scale test launches takes place to verify overall functionality of the vehicle.
The actualization of these is to represent the competition conditions, and one of
these in particular is useful as the full scale demonstration flight before the FRR.
This test is extremely vital to confirming the design, as it requires all respective
components of the vehicle to perform as intended.
Timed assemblies of the launch vehicle ensure that all components integrate and
that the vehicle is ready for a launch within the time limit at competition. Tuning up
the team for timely assemblies reduces the chance for potential error in preparing
the vehicle for launch.
Planning of Integration
The project manager and lead engineer are present at all team meetings and
subsystem meetings. They carry the responsibility of ensuring proper and efficient
communication of the group, such that all necessary subsystems of the vehicle
design integrate successfully according to the plan. Three team meetings per
week are part of the general schedule, which allows the subsystem leads to
present their progress in front of the entire team. These sessions allow the team to
address any design issues, concerns, or questions.
Planning of Operations
All operations relating to this project must have a schedule, procedure, and
checklist to ensure all steps leading to the successful completion of each operation
happens. This ensures efficiency in carrying out a particular operation and allows
checklists and procedures to develop in accordance with safety regulations. All

45
test flights follow the checklist found in Appendix B. Prior to each flight, testing
verification on each component takes place to ensure flight readiness.
The team is divided into several subsystems. Figure 13 is a hierarchy chart of the
team.
Figure 13 - Team Hierarchy
Confidence and Maturity of Design
The vehicle design evolves throughout the design process. Every change seeks to
improve the overall design. Over time, the flaws and failures of the design lose out,
helping the design to mature. The more time analyzing and testing should increase
the opportunities for evolution. The overall maturity of the design should increase
throughout these evolutionary stages. Already, the design has progressed through
several evolutionary phases. The current design is at an intermediate maturity.
While inevitably many design flaws persist, many cease to be and changes
continually mount to improve upon the design. It is necessary to follow through on
planning and execution of necessary maturity/risk reduction efforts throughout the
product life cycle. Figure 14 shows the maturity life cycle of the project.

46
Figure 14 - Project Life Cycle
The Tarleton Aeronautical Team is confident in their design abilities. Although it is
certain the design will evolve, the current design has progress to speak for itself.
Extensive hours formulating the current design have been worth the effort. The
team is confident in the overall design of the vehicle and the ability to achieve the
target goals in the competition. The design chosen for this year’s competition is
simple and efficient, with a clear, modular payload system. The team examines
every subsection in detail for flaws or possible improvements on a theoretical level.
Tests on each component continually reveal further information regarding the
design’s maturity.
Dimensional Drawing
The major sections of the vehicle are represented in Figure 15.
Figure 15 - Dimensional Drawings

47
Electrical Schematics of the Recovery System
The recovery system consists of three main electrical devices: the PerfectFlite
Stratologger recovery altimeter, the Featherweight Raven 3 recovery altimeter,
and the BeeLine GPS. The recovery altimeters each utilize a dedicated nine volt
power supply, and the BeeLine GPS utilizes a five volt power supply. Correct
wiring of the recovery altimeters is crucial to a safe and successful recovery.
Improper wiring could cause inadvertent deployment of the recovery system,
risking injury to people and the vehicle. Figure 16 is a conceptual wiring schematic
of the recovery’s electronic components. There are two setups of the recovery
system electronics; one for the drogue deployment and one for the main.
Furthermore, there are two recovery GPS modules.
Figure 16 - Recovery Electrical Schematic
Figure 17 is a picture of the Featherweight Raven 3 deployment altimeter.

48
Figure 17 - Featherweight Raven 3 Wiring
Mass Statement
The mass summary of the vehicle is in Table 14. Each subsection breaks down
into its respective components in Tables 15 through 17. The mass calculations for
the launch vehicle, subsections, and individual components come from three
methods. The mass of components is retrievable from data sheets when available.
Density of the materials and volume of the structural components help obtain mass
estimates. Where no data is available, logical deductions provide reason towards
the component mass based on similarity to other known components. This allows
for a reasonable level of accuracy and to allow a reserve of three to five pounds for
a possible mass growth. Concluding from the listed mass of 33.5 lbs for the launch
vehicle and the maximum thrust of 437.7 lbf from the propulsion system, the rocket
has a thrust to weight ratio of 13:1. This requires more than 400 lb of additional
mass to prevent the vehicle from launching.
Overall
Subsection Mass (oz) Mass (lb)
Payload 40.58 2.54
Recovery 59.85 3.74
Structure 435.6 27.23
Total Mass (Launch) 536.03 33.50
Total Mass (Apogee) 473.92 29.62
Table 14 Total Mass Summary

49
Payload
Component Quantity Mass (oz) Total Mass (oz)
Baseplate 1 1.11 1.11
Battery 8 1.28 10.24
Circuit Boards 1 6 6
Railing – Main 2 2.8125 5.625
Railing – Support 2 0.262 0.524
Sensors/Electronics 1 13.1 13.1
Servo – Large 1 1.55 1.55
Servo – Small 1 0.67 0.67
Video Camera 1 1.76 1.76
Subtotal 40.579
Table 15 Payload Mass Summary
Recovery
Attachment Hardware 2 3 6
Charges – Drogue 1 3 3
Charges – Main 1 4 4
Deployment Bag – Drogue 1 3 3
Deployment Bag – Main 1 5 5
GPS 2 2 4
Parachute – Drogue 1 2.63 2.63
Parachute – Main 1 11.3 11.3
Recovery Electronics – Drogue 1 5 5
Recovery Electronics – Main 1 5 5
Shock Cord – Drogue 1 4.68 4.68
Shock Cord – Main 1 6.24 6.24
Subtotal 59.85
Table 16 Recovery Mass Summary
Structure

50
Acrylic Payload Section 1 52.3 52.3
Ballast 1 10.92 10.9
Bulkhead 3 3.03 9.09
Bulkhead – Motor 1 6.07 6.07
Bulkhead – Payload 2 32.8 65.6
Center Rings 3 2.01 6.03
Coupler 2 14.3 28.6
Engine Compartment 1 12.9 12.9
Body Tube – Front 1 38.4 38.4
Body Tube – Rear 1 49.4 49.4
Fin 4 5.625 22.5
Motor 1 118 118
Nosecone 1 15.8 15.8
Subtotal 435.6
Table 17 Structure Mass Summary
Recovery System
Deployment of Parachutes
A dual-stage deployment
recovery system is in use,
consisting of the staged
release of a drogue
parachute and a main
parachute. The main
parachute ejects from the top
of the upper body structure,
just below the nose cone.
The drogue parachute ejects
from the drogue parachute
compartment at the front of
the lower body structure.
This staging is in Figure 18.

51
In order to minimize both landing radius and terminal velocity, the drogue
parachute deploys at 5,280 feet, when the vehicle is at apogee. The main
parachute deploys at 500 feet above ground level on descent to ensure proper
final velocity. It is imperative that the drogue parachute deploys at apogee in order
to avoid damage to the rocket body caused by the jarring that would ensue due to
a high speed ejection.
Deployment Altimeters
In order to eliminate the variability of choosing the right delay time and to improve
redundancy, each deployment functions with two altimeters. Each altimeter system
consists of a main altimeter, backup altimeter, and e-match wiring. The main
altimeter is a Featherweight Raven3 and the backup is a PerfectFlite StratoLogger
which is completely independent of the payload electronics.
Each altimeter system is inside of a vented compartment below each parachute
compartment in the vehicle body. Each altimeter has its own dedicated power
supply, a standard 9-volt battery. Each altimeter system mounts vertically on a
0.125-inch-thick, 4-inch-wide, 1-inch-long fiberglass board. One altimeter is on
each side. Each board then epoxies on either end to a 0.125-inch-thick, 5.375-
inch-diameter fiberglass disk. The entire setup bolts to the bulkhead below each
parachute compartment.
The compartments seal from the black powder ejection charges. Each
compartment vents to ambient air pressure in order to acquire proper altitude
readings. There is a porthole drilled from the exterior of the rocket body into each
altimeter compartment, the size of which shall be determined through later
testing.
Item Distributor Product
Unit
Dimensions
Unit
Cost
Number
Total
Cost
Main
Altimeter
Featherweight
Altimeters
Raven3
1.8in. X 0.8in. X
0.55in. X 0.34oz
$155.00 2 $310.00
Loki Research Ozark ARTS
3.75in. long X
1.4in. wide X
2.75 oz
$190.00 2 $380.00
Backup
Altimeter
PerfectFlite Stratologger
2.75in. long X
0.9 in. wide X
0.45 oz
$79.95 2 $159.90
Adept Rocketry ALTS1-50K
0.9in. X 0.65in. X
4.25in. X 4.25oz
$89.00 2 $178.00
Table 18 Deployment Altimeter Trade and Selection

52
The Featherweight altimeter is the main altimeter due to its versatility. It has full
functionality regardless of positioning, has visible and audible readout of individual
channel continuity and battery voltage, allows for user calibration of the
accelerometer rather than presets, can record up to eight minutes of high-rate data
plus an additional 45 minutes per flight, and has a downloadable interface program
which is easy to read. The audible readout function deactivates manually prior to
launch. Table 18 shows that the Featherweight is the choice for the project.
Though features are comparable, physical dimensions are not. Optimal
engineering efficiency comes through employment of the Featherweight. The cost
of the Featherweight is less than that of the Ozark.
The PerfectFlite altimeter is the backup altimeter due to its high level of reliability.
False triggering is not a problem for gusts of wind up to 100 miles per hour. The
precision sensor and 24-bit analog-to-digital converter (ADC) allow for 99.9
percent accurate altitude readings, and the selectable apogee delay for dual
setups prevents overpressure from simultaneous charge firing. As demonstrated in
Table 18, the PerfectFlite proves superior to the ALTS1-50K. While features
compare well, the ALTS1-50K is not the best choice concerning compartment
space capacity or cost.
Additionally, each altimeter system has an externally-accessible magnetic arming
switch capable of being locked in the “on” position for launch. The arming switch
dedicated to the dual altimeters, which control main parachute deployment, are at
five feet, eight inches above the base of the launch vehicle. Those dedicated to the
dual altimeters which control drogue parachute deployment are two feet above the
base of the launch vehicle.
Ejection Charges
In order to ensure separation and ejection of the proper parachute at the proper
time, each altimeter is set to light a one-foot low-current Daveyfire N28BR electric
match. There are holes between the lower bulkhead and altimeter compartment to
allow the lead on each e-match through; the holes are shut with epoxy to seal the
chamber from the other chambers in the vehicle body. The basic construction of
Figure 19 Daveyfire Electric Match

53
an electric match, or e-match, is in Figure 19.
The e-match selection is indicated in Table 19. While the Daveyfire is more costly
than the QuickBurst or RocketFlite, it is the only fully-assembled option. The
QuickBurst and RocketFlite are less costly options, but each requires manual
assembly. Improper assembly of an e-match could result in electric shock and
premature ignition of ejection charges.
Distributor Item
Unit
Dimensions
Pre-
Assembled
Unit
Cost
Coast
Rocketry
Daveyfire N28BR 1 ft long X 1 Yes $2.95
QuickBurst
QuickBurst E-Match
Kit
1 ft long X 20 No $32.00
RocketFlite MF-12 1 ft long X 12 No $9.95
Table 19 Electric Match Trade & Selection
Assuming that the entire mass of each charge is burns and converts into a gas,
the basic Ideal Gas Law is used,
( )
With , and for the volume of the cylinder,
( ) ( )
( ) ( )
( )
For a compartment length of for the drogue ejection,

54
For a compartment length of for the main ejection,
Each black powder ejection charge consists of a well of black powder contained in
a plastic charge tube with the shroud of an e-match immersed in the well.
Global Positioning System
A BeeLine GPS is the selection for recovering each component upon landing in
the event that tethering separation on descent occurs or visual contact is lost.
Each GPS is in a 2.5 inch sub-compartment of each parachute compartment. In
order to achieve this, each GPS mounts to a 0.125 inch thick, 2.25 inch long, four
inch wide sheet of fiberglass with epoxy, then on either end to a 0.125 inch thick
and 5.375 inch wide fiberglass disk which is inserted below the lower bulkhead of
each altimeter compartment. This should shield the devices from parachute
ejection, black powder
charge ignition, and fuel ejection.
A fine copper wire mesh lines the internal surface of the drogue altimeter housing,
which is separate from the GPS sub-compartment via bulkhead, such that these
altimeters are shielded from radio frequencies in order to prevent inadvertent
excitation. The holes in this mesh must be significantly smaller than the
wavelength of the interfering radio frequencies so that the enclosure does not
ineffectively approximate an unbroken conducting surface. The BeeLine GPS
operates in the range 420 to 450 mHz. The XBee operates at 900 mHz. A pure
copper mesh fabric with electromagnetic frequency blocking effectiveness in the
range 900 to 420 mHz is the team’s choice to line each compartment containing a
BeeLine GPS.
A corresponding ground receiver is in the ground station. Each BeeLine package
includes a fully integrated RF transmitter, GPS and RF antennas, GPS Module,
and battery. Altogether, these devices simultaneously transmit latitude, longitude,
altitude, course, and speed. These quantities are analyzable after each flight in
order to aid in continued optimization.
The BeeLine GPS has been chosen for its small size, reasonable cost,
transmission range at up to 20 miles line of sight, frequent usage in high-powered
model rocketry, use of standard decoding hardware (automatic packet reporting
system, or APRS), and operation frequency on any frequency in the 70-centimeter
amateur radio band. Additionally, the BeeLine is the only consistently commercially
available fully integrated GPS system for model rockets. Its measurements of
course and speed allow for real-time calculation of landing distance and terminal
kinetic energy. These measurements also serve as a check for the altimeters. The
mounting precautions, the 8 hour battery life of the Lithium-Poly battery, the non-
volatile flight-data memory storage (3 hours at 1 Hertz), the user-programmable

55
transmission rates, and output power will ensure that the devices remain fully
functional during the course of the flight.
Recovery Testing
Static testing on each set of altimeters with e-matches ensures the reliability of the
electronic system. The possibility of delaying the signal from the backup altimeter
by up to two seconds with respect to the main altimeter is under consideration.
This could help to ensure that separation does occur should black powder well
leak, humidity become a problem, or pressure conditions prevent a sufficiently
powerful force from the main charge.
Once static testing on the altimeter-e-match systems is complete, ground testing of
the system with the charge wells in an empty replica of the launch vehicle is
conducted under the guidance of the team mentor in the presence of the team
safety officer. This enables any sizing adjustments which may be necessary to
ensure vehicle separation are addressed prior to conducting a full-scale launch.
The final testing phase of the recovery system components includes at least one
subscale single-event flight, at least one subscale dual-event flight, and lastly at
least one full-scale dual-event flight. The purpose of the subscale single-event
flight is to allow for proper understanding of the parachute components in
combination with the electronic charge system. Once the parachute components
perform in real a situation, a subscale dual-event flight will serve to mimic the
competition flight in a more contained way. Finally, a series of full-scale dual-event
flights confirms expectations for the competition flight. The cycle of these tests is in
Figure 62, but specific data to the recovery system is in Table 20, where one
denotes static testing, two denotes sub-scale, and three denotes full-scale.
Figure - Recovery Testing Dates
Month Date Stage Verification Event
October 27 2 Dual Deployment Test Launch
November 12 1 Parts Ordered for Prototyping
17 3 Dual Deployment Launch
30 1 End of Lab Prototyping
December 1 2 Low-Altitude Flight
3 3 Motor Assembly
5 1 End of Programming
8 3 Full Scale Launch
22 2 Low Altitude Flight
January 2 2 End of Field Testing
5 3 Test Launch

56
12 3 Alternative Launch
19 1 Static Motor Test
26 3 Low Altitude Full Force of Impact Test Launch
Table 20 Recovery Testing Dates
Anemometer
In order to anticipate the effectiveness of the recovery system prior to flight, a
compact rotary-fan digital anemometer is incorporated into the ground station. As
indicated in the below Table 21, the SpeedTech WindMate-300 (WM-300) is the
team’s choice for its low price in comparison to newer, similar products. Like the
more costly options listed, the WM-300 provides a digital readout of wind speed,
wind direction, humidity, and pressure. It is also resistant to water damage and has
a threaded base which may be mounted to a tripod at the ground station. Mounting
the anemometer to a tripod maintains stability of the device to ensure accurate
readings.
Component Distributor Item
Unit
Dimensions
Wind
Speed
and
Direction
Water
Proof
Mount
for
Tripod
Unit
Cost
4500 Pocket
Weather
Tracker
Ambient
Weather
Kestrel
4500
5in. X 1.8in.
X 1.1in.
Yes Yes Yes $299.00
WindMate
Anemometer
Weather
Shack
Speed
Tech
WM-
350
5.5in. X
1.75in. X
0.75in.
Yes Yes Yes $229.95
WindMate
Anemometer
Weather
Shack
Speed
Tech
WM-
300
5.5in. X
1.75in. X
0.75in.
Yes Yes Yes $154.95
Table 21 Anemometer Trade & Selection
Parachute Size Calculations
Assuming use of a Cesaroni L1720-WT-P motor, the total vehicle launch weight is
estimated to be 29.62 pounds, as seen in Table 14. The team takes into
consideration the addition of up to 10 percent ballast, as well as the fuel
compartment being empty by deployment of the drogue parachute, and estimates
the vehicle weight at 29.62 pounds. While the weight estimate may change, the
process of the following calculations will not.

57
The following calculations give the maximum descent rate upon landing for the
vehicle to have a kinetic energy of less than 75 foot pound force.
( )
√
( )
Therefore for the vehicle to land with a kinetic energy of less than 75 foot pound
force, based on a weight of 29.18 pounds, the descent rate must be less than
12.8656 feet per second. Using this result, it is possible to calculate the minimum
diameter of the main parachute as follows:
( )
( )

58
∑
.
( )
( ) ( )
√
√
Therefore, for the vehicle to land with a kinetic energy of less than 75 foot pound
force, the main parachute must be at least 9.965 feet in diameter. Due to
commercial availability, the main parachute diameter is 10 feet. The following
calculates the exact descent rate of the vehicle based on the 10 foot diameter
main parachute.

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