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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 ...................................................................................... 79
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 ......................................................................................................... 86
Environment ............................................................................................................... 88
Environmental effects of the project........................................................................ 88
Environmental effect on the project ........................................................................ 89
IV) Payload Criteria ....................................................................................................... 91
System Level Review ............................................................................................. 91
Required Subsystems ................................................................................................ 95
Atmospheric Data Gathering .................................................................................. 95
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 .................................................................................... 115
Verification Plan ....................................................................................................... 116
Preliminary Integration Plan ..................................................................................... 118
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 ............................................................................................................ 141
Testing Budget...................................................................................................... 141
Outreach Budget................................................................................................... 142
Travel Budget ....................................................................................................... 142
Budget Summary .................................................................................................. 142
Funding Plan ............................................................................................................ 143
Timeline ................................................................................................................... 143
Gantt Timeline ...................................................................................................... 144
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 ................................................................................. 55
Table 21 Anemometer Trade & Selection ..................................................................... 56
Table 22 Parachute Diameters ...................................................................................... 61
Table 23 Main Parachute Trade and Selection ............................................................. 61
Table 24 Drogue Parachute Trade and Selection ......................................................... 62
Table 25 Parachute Protection Materials ...................................................................... 62
Table 26 Shock Chord Trade and Selection .................................................................. 63
Table 27 Launch Day Recovery System Budget ........................................................... 65
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 .......................................................................................... 118
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
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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 Aeronautical Team
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
Outer Diameter
Mass
Motor
Selection
Recovery
Drogue
Main
108 inches
5.525 inches
33.5 pounds
Cesaroni L1720-WT-P
24” Nylon Parachute, Apogee Deployment
120” Nylon Parachute, 500 foot AGL Deployment
Primary Featherweight Raven 3 Altimeter,
Backup PerfectFlite Stratologger Altimeter, and Beeline GPS Tracking
Table 1 Launch Vehicle Summary
Avionics
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
To provide shielding from wireless transmitting devices
compartments
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
Number of pyranometers from 2 to 4
For efficiency of design
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
Two, 90 degree opposing boards for solar
irradiance and UV sensors
For structural support
Solar irradiance and UV sensors can gather data at 90
degree intervals around the rocket vehicle
Handheld Yagi directional antenna will be used to
Development of an automated tracking system for the
receive all transmitted data from the payload
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
To increase quality of Educational outreach
aspect to the entire project
To increase quality of public outreach and
project sustainability
Table 5 Project Plan Changes
Educational Outreach Plan has been further developed
Community Outreach events added
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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.
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
Material
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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2012 - 2013 USLI Preliminary Design Review
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
Inability of external
inspection, externally
mounted sensors,
inaccessibility to
internal systems
Durability, weight
Inability
ofand
external
Table
7 Pros
Cons of Payload Options
Cons
inspection, externally
mounted sensors,
coupling system
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
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2012 - 2013 USLI Preliminary Design Review
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.
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2012 - 2013 USLI Preliminary Design Review
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:
Selected Option
Alternative Option 1
Figure 7 - Fin Design Options
Alternative Option 2
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2012 - 2013 USLI Preliminary Design Review
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
Figure 8 Selected Fin Design
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
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2012 - 2013 USLI Preliminary Design Review
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)
Cesaroni
L1720WT-P
5345
75.9
Aerotech
L1390G
5618
65.6
Cesaroni
L1090SSP
6479
76.8
Total
Impulse
Max.
Velocity
(ft./s)
Average
Thrust
830.9lbf
s
394.3lbf
738
(3696Ns
(1754N)
)
887.8lbf
s
308.9lbf
723
(3949Ns
(1374N)
)
1082lbfs
246.6lbf
(4815Ns
733
(1097N)
)
Table 8 Motor Trade and Selection
Burn
Time
(s)
Thrust
to
Weight
Ratio
Availability
/
Cost
2.15
11.8
High/
$170.96
2.65
9.2
Medium/
$209.99
4.4
7.4
Medium/
$346.95
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
𝑇
1754𝑁
motor the thrust to weight ratio is achievable by 𝑊 = 15.2𝑘𝑔∗9.8𝑚/𝑠2 = 11.8 : 1, which
exceeds the suggested ratio of 5 : 1.
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2012 - 2013 USLI Preliminary Design Review
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 highpowered rocket, the thrust to weight ratio is given by
T
W
=
Average Engine Thrust,in Newtons
.
(Rocket Launch Mass,in kilograms)∗(Acceleration of Gravity,in meters per second squared)
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2012 - 2013 USLI Preliminary Design Review
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
𝑇
1374𝑁
ratio is given by 𝑊 = 15.8𝑘𝑔∗9.8𝑚/𝑠2 = 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
𝑇
1097𝑁
weight ratio is given by 𝑊 = 17.3𝑘𝑔∗9.8𝑚/𝑠2 = 6.47 : 1, which exceeds the suggested
ratio of 5 : 1.
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2012 - 2013 USLI Preliminary Design Review
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
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2012 - 2013 USLI Preliminary Design Review
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:
𝐾𝑖𝑛𝑒𝑡𝑖𝑐 𝐸𝑛𝑒𝑟𝑔𝑦 = 𝐾𝐸 =
1
𝑚𝑣 2
2
∑ -𝐹𝑧 = 𝐷𝑟𝑎𝑔 − 𝑊𝑒𝑖𝑔ℎ𝑡 = 𝑚𝑔 = 0
1
𝐷𝑟𝑎𝑔 = 𝑊𝑒𝑖𝑔ℎ𝑡 = 𝑝𝑣 2 𝑆𝐶𝑑
2
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.
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2012 - 2013 USLI Preliminary Design Review
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.
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2012 - 2013 USLI Preliminary Design Review
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
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
1.2.2
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2012 - 2013 USLI Preliminary Design Review
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
Tarleton Aeronautical Team 35
2012 - 2013 USLI Preliminary Design Review
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
Tarleton Aeronautical Team 36
2012 - 2013 USLI Preliminary Design Review
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
N/A
N/A
The schedule and budget
plan for the full scale
motor to be flown.
Schedule
1.15.2.1
1.15.2.3
1.15.3
Unmanned aerial vehicles, and/or
recovery systems that control
flight path of vehicle, will fly as
designed during full scale
demonstration flight
Full scale motor does not have to
be flown during full scale test
flight
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 fullscale 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
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
1.16
Tarleton Aeronautical Team 37
2012 - 2013 USLI Preliminary Design Review
Table 9 Vehicle Verification Table
Table 10 is the verification table for the recovery system.
Requirement
(SOW)
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
Recovery System
Requirement
Tarleton Aeronautical Team 38
2012 - 2013 USLI Preliminary Design Review
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 onboard 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
Tarleton Aeronautical Team 39
2012 - 2013 USLI Preliminary Design Review
2.12.2
Recovery electronics shall be
shielded from all on-board
transmitting devices
Altimeter compartments
are shielded by intenal
copper mesh lining
Inspection,Testing
2.12.3
Recovery electronics shall be
shielded from any magnetic
waves generated by onboard devices
Altimeter compartments
are shielded by intenal
copper mesh lining
Inspection,
Testing
2.12.4
Recovery electronics shall be
shielded from any on-board
device that could adversely
affect proper operation
Altimeter compartments
are shielded by intenal
copper mesh lining;
Each altimeter bay has
a dedicated and
separate compartment
Inspection,Testing
2.13
Recovery system shall use
commercially available lowcurrent e matches for ignition
of ejection charges
Davyfire N28BR ematches 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
Safety
Unable to
Obtain Flight
Waivers
Disregarding
Safety Plan
Medium
Medium
Low
High
Consequence
Delay in Test
Launches
Harm to
Participants
Mitigation
Schedule Test
Flight at
Official Test
Launch Sites
Strict Safety
Plan
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2012 - 2013 USLI Preliminary Design Review
Enforcement
Structure
Propulsion
Educational
Engagement
Electronics
Recovery
Ignorance of
Legal Rules and
Regulations
Destruction of
Parts During
Testing
Improper Test
Vehicle
Assembly
Inconsistent
Motor
Construction
Incorrect
Motor Mount
Assembly
Poor Quality
Presentation
Scheduling
Conflict With
The Schools
Recurring Legal
Education
Low
Legal
Repercussions
Reordering of
Parts, Delay in
Project
Medium
Medium
Unpredictable
Performance
Low
High
Assembly
Checklist
Thorough
Inspection and
Analysis
Low
High
Low
High
Medium
High
Low
High
High
Unpredictable
Performance
Possible Harm
to Vehicle and
Participants
Less Effective
Outreach
Low
Medium
Low
Low
Power Budget
Miscalculation
Low
High
Incorrect Black
Powder Rating
Low
High
Parachute Size
Miscalculation
Low
Low
Cancellation of
Events
Delay to
Project,
Ineffective
Payload
Harm to Parts,
Ineffective
Payload
Harm to Parts,
Ineffective
Payload
Harm to
Vehicle and
Participants
Violation of
Competition
Rules
Medium
Lack of
Funding and
Support
Medium
Medium
Lack of
Funding and
Support
Delay in
Faulty
Components
Incorrect
Wiring
Configurations
Public
Relations
Lack of Project
Exposure
Management
Harmful
Representation
of Team
Low
Inadequate
High
Low
Inventory Extra
Parts
Assembly
Checklist
Rehearsed
Presentation
Communicate
With Schools
Ordering
Duplicates,
Testing
Thorough
Research and
Design
Redundant
Calculations
Extremely
Thorough
Testing
Redundant
Calculations
Constant
Outflow of
Updates and
Information
Proper Ethics
and
Professionalis
m
Weekly Team
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2012 - 2013 USLI Preliminary Design Review
Communicatio
n
Scheduling
Conflicts
Software
University
Administration
Delays
Loss of Source
Code
Incorrect
Algorithm
Design
Incorrect
Datasheets
Medium
Medium
Medium
Low
Medium
High
Low
Medium
Low
Low
Project, Project
Failure, Design
Flaws
Delay in
Project
Meetings,
Email Minutes
Delay in
Project/Fundin
g
Rewrite
Software
Maintain
Scheduling
Proper
Communicatio
n and
Scheduling
Backup Source
Code
Ineffective
Payload Logic
Confusion,
Delay in
Project
Redundant
Calculations
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.
Tarleton Aeronautical Team 42
2012 - 2013 USLI Preliminary Design Review
Figure 12 - Risk Plot
Severity #
Severity
Probabilty
Severity #
LOW
Severity
Probabilty
Severity #
Medium
Severity
High
20 Delay in Deliveries
90
60 Communication Failure
33 Destruction of Parts During Testing
85
1 Incorrect Wiring Configurations
50
3 Parachute Size Miscalculation
10
39 Faulty Components
30 University Administration Delays
40
33 Lack of Project Exposure
20
34 Harmful Representation of Team
10
50 Inadequate Communication
66
55 Scheduling Conflicts
75
67 Teammates Disregarding Safety Plan
Ignorance of Legal Rules and
99 Regulations
35 Incorrect Algorithm Design
55
72 Inconsistent Motor Construction
57 Environment Prevents Recovery
25
75 Loss of Source Code
52 Poor Quality Presentation
10
33 Poor Weather
29 Scheduling Conflict With The Schools
12
66 Burn Ban in Effect
5 Incorrect Datasheets
5
Probability
10
90 Inadequate Personnel
25
66 Unable to Obtain Flight Waivers
0
80 Manufacturing Issues
20
45 Improper Test Vehicle Assembly
15
99 NAR/TRA Violations
5
4
98 Damage of Property
15
70 OSHA Violations
10
100 Personal Injury
1
42 Power Budget Miscalculation
Table 12 Project Risks
1
2
3
22
50
18.5
25
92 Loss of Rocket Lab
Loss of Low-Altitude Test Launch
93 Facility in Glen Rose, TX
Loss of High-Altitude Test Launch
94 Facility in Cross Plains, TX
5
38 Loss of Science Building
0
5
5
39 Engineering Building
0
40 Loss of HCI facilities
12
Tarleton Aeronautical Team 43
2012 - 2013 USLI Preliminary Design Review
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
Structure Testing
Lab Prototyping
Low Altitude flight
Dual deployment
Force of impact
Full scale launch
Timed final assembly
Subsystem
Structure
Structure, Recovery, Integration
Structure, Recovery
Recovery
Structure, Recovery, Integration
Vehicle, Recovery, Integration
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
Tarleton Aeronautical Team 44
2012 - 2013 USLI Preliminary Design Review
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
Tarleton Aeronautical Team 45
2012 - 2013 USLI Preliminary Design Review
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.
Tarleton Aeronautical Team 46
2012 - 2013 USLI Preliminary Design Review
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
Tarleton Aeronautical Team 47
2012 - 2013 USLI Preliminary Design Review
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.
Tarleton Aeronautical Team 48
2012 - 2013 USLI Preliminary Design Review
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
Payload
Recovery
Structure
Total Mass (Launch)
Total Mass (Apogee)
Mass (oz)
40.58
59.85
435.6
536.03
473.92
Table 14 Total Mass Summary
Mass (lb)
2.54
3.74
27.23
33.50
29.62
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2012 - 2013 USLI Preliminary Design Review
Payload
Component
Baseplate
Battery
Circuit Boards
Railing – Main
Railing – Support
Sensors/Electronics
Servo – Large
Servo – Small
Video Camera
Quantity
1
8
1
2
2
1
1
1
1
Mass (oz)
1.11
1.28
6
2.8125
0.262
13.1
1.55
0.67
1.76
Subtotal
Total Mass (oz)
1.11
10.24
6
5.625
0.524
13.1
1.55
0.67
1.76
40.579
Table 15 Payload Mass Summary
Recovery
Component
Attachment Hardware
Charges – Drogue
Charges – Main
Deployment Bag – Drogue
Deployment Bag – Main
GPS
Parachute – Drogue
Parachute – Main
Recovery Electronics – Drogue
Recovery Electronics – Main
Shock Cord – Drogue
Shock Cord – Main
Quantity
2
1
1
1
1
2
1
1
1
1
1
1
Mass (oz)
3
3
4
3
5
2
2.63
11.3
5
5
4.68
6.24
Subtotal
Table 16 Recovery Mass Summary
Total Mass (oz)
6
3
4
3
5
4
2.63
11.3
5
5
4.68
6.24
59.85
Tarleton Aeronautical Team 50
2012 - 2013 USLI Preliminary Design Review
Structure
Component
Acrylic Payload Section
Quantity
1
Ballast
Bulkhead
Bulkhead – Motor
Bulkhead – Payload
Center Rings
Coupler
Engine Compartment
Body Tube – Front
Body Tube – Rear
Fin
Motor
Nosecone
Subtotal
Mass (oz)
52.3
1
10.92
10.9
3
1
2
3
2
1
1
1
4
1
3.03
6.07
32.8
2.01
14.3
12.9
38.4
49.4
5.625
118
9.09
6.07
65.6
6.03
28.6
12.9
38.4
49.4
22.5
118
1
15.8
15.8
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.
Total Mass (oz)
52.3
Tarleton Aeronautical Team 51
2012 - 2013 USLI Preliminary Design Review
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, 5.375inch-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
Main
Altimeter
Backup
Altimeter
Distributor
Product
Unit
Dimensions
Unit
Cost
Number
Total
Cost
Featherweight
Altimeters
Raven3
1.8in. X 0.8in. X
0.55in. X 0.34oz
$155.00
2
$310.00
Loki Research
Ozark ARTS
$190.00
2
$380.00
PerfectFlite
Stratologger
$79.95
2
$159.90
Adept Rocketry
ALTS1-50K
$89.00
2
$178.00
3.75in. long X
1.4in. wide X
2.75 oz
2.75in. long X
0.9 in. wide X
0.45 oz
0.9in. X 0.65in. X
4.25in. X 4.25oz
Table 18 Deployment Altimeter Trade and Selection
Tarleton Aeronautical Team 52
2012 - 2013 USLI Preliminary Design Review
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
Figure 19 Daveyfire Electric Match
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
Tarleton Aeronautical Team 53
2012 - 2013 USLI Preliminary Design Review
chamber from the other chambers in the vehicle body. The basic construction of
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.
Item
Unit
Dimensions
PreAssembled
Unit
Cost
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
Distributor
Coast
Rocketry
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,
𝑃𝑉 = 𝑁𝑅𝑇
𝑃 = 𝑑𝑒𝑠𝑖𝑔𝑛 𝑝𝑟𝑒𝑠𝑠𝑢𝑟𝑒 𝑖𝑛 𝑝𝑜𝑢𝑛𝑑𝑠 𝑝𝑒𝑟 𝑠𝑞𝑢𝑎𝑟𝑒 𝑖𝑛𝑐ℎ
𝑉 = 𝑣𝑜𝑙𝑢𝑚𝑒 𝑜𝑓 𝑐𝑦𝑙𝑖𝑛𝑑𝑟𝑖𝑐𝑎𝑙 ℎ𝑜𝑢𝑠𝑖𝑛𝑔 𝑐𝑜𝑚𝑝𝑎𝑟𝑡𝑚𝑒𝑛𝑡
𝑁 = 𝑚𝑎𝑠𝑠 𝑜𝑓 𝑝𝑜𝑤𝑑𝑒𝑟 𝑖𝑛 𝑝𝑜𝑢𝑛𝑑𝑠
𝑅 = 𝑢𝑛𝑖𝑣𝑒𝑟𝑠𝑎𝑙 𝑔𝑎𝑠 𝑐𝑜𝑛𝑠𝑡𝑎𝑛𝑡 (𝑖𝑛 • 𝑙𝑏𝑓 /𝑙𝑏𝑚)
𝑇 = 𝑖𝑛𝑖𝑡𝑖𝑎𝑙 𝑐𝑜𝑚𝑏𝑢𝑠𝑡𝑖𝑜𝑛 𝑡𝑒𝑚𝑝𝑒𝑟𝑎𝑡𝑢𝑟𝑒 𝑜𝑓 𝑝𝑜𝑤𝑑𝑒𝑟 𝑖𝑛 𝑅𝑎𝑛𝑘𝑖𝑛𝑒
With 𝑃 = 15 𝑙𝑏/𝑖𝑛2 , and 𝑉 = 0.25𝜋𝐷2 𝐿 for the volume of the cylinder,
𝑁=
𝑃𝑉
𝑅𝑇
𝑙𝑏𝑓
) (7.223𝜋𝐿 𝑖𝑛3 )
𝑖𝑛2
𝑁=
𝑖𝑛 ∗ 𝑙𝑏𝑓
(266
) (3307 𝑅)
𝑙𝑏𝑚 ∗ 𝑅
(15
𝑁 = (3.8694 ∗ 10−4 ) ∗ 𝐿 𝑙𝑏𝑚
For a compartment length of 𝐿 = 28 𝑖𝑛 for the drogue ejection, 𝑁 = 0.0108 𝑙𝑏𝑚
For a compartment length of 𝐿 = 36 𝑖𝑛 for the main ejection, 𝑁 = 0.01393 𝑙𝑏𝑚
Tarleton Aeronautical Team 54
2012 - 2013 USLI Preliminary Design Review
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 nonvolatile flight-data memory storage (3 hours at 1 Hertz), the user-programmable
transmission rates, and output power will ensure that the devices remain fully
functional during the course of the flight.
Tarleton Aeronautical Team 55
2012 - 2013 USLI Preliminary Design Review
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.
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
1
2
Low-Altitude Flight
3
3
Motor Assembly
5
1
End of Programming
8
3
Full Scale Launch
22
2
Low Altitude Flight
2
2
End of Field Testing
5
3
Test Launch
12
3
Alternative Launch
19
1
Static Motor Test
3
Low Altitude Full Force of Impact Test Launch
December
January
26
Table 20 Recovery Testing Dates
Tarleton Aeronautical Team 56
2012 - 2013 USLI Preliminary Design Review
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.
Wind
Speed
and
Direction
Water
Proof
Mount
for
Tripod
Unit
Cost
Component
Distributor
Item
Unit
Dimensions
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
WM350
5.5in. X
1.75in. X
0.75in.
Yes
Yes
Yes
$229.95
WindMate
Anemometer
Weather
Shack
Speed
Tech
WM300
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.
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.
Tarleton Aeronautical Team 57
2012 - 2013 USLI Preliminary Design Review
𝑣 = 𝑑𝑒𝑠𝑐𝑒𝑛𝑡 𝑟𝑎𝑡𝑒
𝑤𝑒𝑖𝑔ℎ𝑡 = 𝑚𝑔 = 29.62𝑙𝑏𝑠
𝑔 = 𝑔𝑟𝑎𝑣𝑖𝑡𝑦 = 32.2𝑓𝑡/𝑠 2
𝐾𝑖𝑛𝑒𝑡𝑖𝑐 𝐸𝑛𝑒𝑟𝑔𝑦 = 𝐾𝐸 =
1
𝑚𝑣 2
2
𝐺𝑖𝑣𝑒𝑛 𝐾𝐸 < 75𝑓𝑡 ∙ 𝑙𝑏𝑓
29.62𝑙𝑏𝑠 ∙ 𝑣 2
< 75𝑓𝑡 ∙ 𝑙𝑏𝑓
2 ∙ (32.2𝑓𝑡/𝑠 2 )
𝑆𝑜𝑙𝑣𝑒 𝑓𝑜𝑟 𝑑𝑒𝑠𝑐𝑒𝑛𝑡 𝑟𝑎𝑡𝑒.
2 ∙ (32.2𝑓𝑡/𝑠 2 ) ∙ 75𝑓𝑡 ∙ 𝑙𝑏𝑓
𝑣<√
29.62𝑙𝑏𝑠
𝑙𝑎𝑛𝑑𝑖𝑛𝑔 𝑑𝑒𝑠𝑐𝑒𝑛𝑡 𝑟𝑎𝑡𝑒 < 12.7697𝑓𝑡/𝑠
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:
𝐴𝑠𝑠𝑢𝑚𝑒 𝑡ℎ𝑒 𝑣𝑒ℎ𝑖𝑐𝑙𝑒 𝑖𝑠 𝑑𝑒𝑠𝑐𝑒𝑛𝑑𝑖𝑛𝑔 𝑎𝑡 𝑎 𝑐𝑜𝑛𝑠𝑡𝑎𝑛𝑡 𝑠𝑝𝑒𝑒𝑑 (𝑠𝑡𝑒𝑎𝑑𝑦 𝑠𝑡𝑎𝑡𝑒).
𝐴𝑠𝑠𝑢𝑚𝑒 𝑡ℎ𝑒 𝑣𝑒ℎ𝑖𝑐𝑙𝑒 𝑠𝑖𝑚𝑝𝑙𝑦 𝑚𝑜𝑣𝑒𝑠 𝑑𝑜𝑤𝑛𝑤𝑎𝑟𝑑 (𝑐𝑜𝑛𝑠𝑡𝑟𝑎𝑖𝑛𝑒𝑑 𝑡𝑜 𝑧 − 𝑎𝑥𝑖𝑠).
𝐷 = 𝑑𝑟𝑎𝑔;
𝑊 = 𝑤𝑒𝑖𝑔ℎ𝑡;
∑ -𝐹𝑧 = 𝐷 − 𝑊 = 𝑚𝑔 = 0
𝑝 = 𝑑𝑒𝑛𝑠𝑖𝑡𝑦 𝑜𝑓 𝑎𝑖𝑟 =
0.075𝑙𝑏
;
𝑓𝑡 3
𝑆 = 𝑠𝑢𝑟𝑓𝑎𝑐𝑒 𝑎𝑟𝑒𝑎 𝑜𝑓 𝑡ℎ𝑒 𝑝𝑎𝑟𝑎𝑐ℎ𝑢𝑡𝑒
𝐶𝑑 = 𝑐𝑜𝑒𝑓𝑓𝑖𝑐𝑖𝑒𝑛𝑡 𝑜𝑓 𝑑𝑟𝑎𝑔 ≈ 2
𝐷=𝑊=
1 2
𝑝𝑣 𝑆𝐶𝑑
2
𝑆𝑜𝑙𝑣𝑒 𝑓𝑜𝑟 𝑠𝑢𝑟𝑓𝑎𝑐𝑒 𝑎𝑟𝑒𝑎 𝑜𝑓 𝑡ℎ𝑒 𝑝𝑎𝑟𝑎𝑐ℎ𝑢𝑡𝑒.
Tarleton Aeronautical Team 58
2012 - 2013 USLI Preliminary Design Review
𝑆=
𝑆=
2𝑊
𝑝𝑣 2 𝐶𝑑
2 ∙ (29.62𝑏𝑠)
32.2𝑙𝑏
∙
3
2
(0.075𝑙𝑏/𝑓𝑡 ) ∙ (12.7697𝑓𝑡/𝑠) ∙ 2 1𝑙𝑏𝑓𝑠 2 /𝑓𝑡
𝑆 = 77.99𝑓𝑡 2
𝐴 = 𝑎𝑟𝑒𝑎 𝑜𝑓 𝑎 𝑐𝑖𝑟𝑐𝑙𝑒
𝐴 = 𝜋𝑟 2
𝑆𝑜𝑙𝑣𝑒 𝑓𝑜𝑟 𝑑𝑖𝑎𝑚𝑒𝑡𝑒𝑟.
𝑑𝑖𝑎𝑚𝑒𝑡𝑒𝑟 = 2 ∙ √
=2∙√
𝐴
𝜋
77.99𝑓𝑡 2
𝜋
𝑚𝑖𝑛𝑖𝑚𝑢𝑚 𝑑𝑖𝑎𝑚𝑒𝑡𝑒𝑟 𝑜𝑓 𝑡ℎ𝑒 𝑚𝑎𝑖𝑛 𝑝𝑎𝑟𝑎𝑐ℎ𝑢𝑡𝑒 = 9.965𝑓𝑡
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.
𝑆=
2𝑊
𝑝𝑣 2 𝐶𝑑
𝑆𝑜𝑙𝑣𝑒 𝑓𝑜𝑟 𝑑𝑒𝑠𝑐𝑒𝑛𝑡 𝑟𝑎𝑡𝑒.
𝑣=√
= √
2𝑊
𝑝𝐶𝑑 𝑆
2 ∙ 29.62𝑙𝑏
32.2𝑙𝑏
∙
3
2
(0.075𝑙𝑏/𝑓𝑡 ) ∙ 2 ∙ 78.5398𝑓𝑡 1𝑙𝑏𝑓𝑠 2 /𝑓𝑡
𝑑𝑒𝑠𝑐𝑒𝑛𝑡 𝑟𝑎𝑡𝑒 = 12.7246𝑓𝑡/𝑠
Tarleton Aeronautical Team 59
2012 - 2013 USLI Preliminary Design Review
The calculated descent rate of the vehicle after main parachute deployment at 500
feet is 12.7246 feet per second. With this figure it possible to calculate the total
drift of the vehicle after main parachute deployment until landing. The following
math calculates the drift of the vehicle from main parachute deployment until
landing assuming a 15 mile per hour (22 feet per second) horizontal wind velocity.
𝑎 = 𝑐ℎ𝑎𝑛𝑔𝑒 𝑖𝑛 𝑎𝑙𝑡𝑖𝑡𝑢𝑑𝑒 = 500𝑓𝑡;
𝑣 = 𝑑𝑒𝑠𝑐𝑒𝑛𝑡 𝑟𝑎𝑡𝑒
𝑤ℎ = ℎ𝑜𝑟𝑖𝑧𝑜𝑛𝑡𝑎𝑙 𝑤𝑖𝑛𝑑 𝑣𝑒𝑙𝑜𝑐𝑖𝑡𝑦 = 22𝑓𝑡/𝑠
𝐷𝑟𝑖𝑓𝑡 =
=
𝑎
∙𝑤
𝑣 ℎ
500𝑓𝑡
∙ 22𝑓𝑡/𝑠
12.7246𝑓𝑡/𝑠
𝐷𝑟𝑖𝑓𝑡 = 864.4673𝑓𝑡
With a 15 mile per hour wind, the vehicle will drift approximately 864 feet after
main deployment at an altitude of 500 feet, until landing. Since the maximum
allowable drift is 2500 feet, the drift between apogee and the main parachute
deployment can only be 1636 feet, 2500 minus 864 feet. The following calculations
give the minimum descent rate for the vehicle while the drogue parachute is
deployed, taking into account a drift of less than 1636 feet.
𝐴𝑠𝑠𝑢𝑚𝑒 𝑎𝑝𝑜𝑔𝑒𝑒 𝑜𝑓 5400 𝑓𝑒𝑒𝑡.
𝐷𝑟𝑖𝑓𝑡 =
𝑎
∙𝑤
𝑣 ℎ
𝑆𝑜𝑙𝑣𝑒 𝑓𝑜𝑟 𝑑𝑒𝑠𝑐𝑒𝑛𝑡 𝑟𝑎𝑡𝑒.
𝑣=
=
𝑎 ∙ 𝑤ℎ
𝐷𝑟𝑖𝑓𝑡
4900𝑓𝑡 ∙ 22𝑓𝑡/𝑠
1636𝑓𝑡
𝑚𝑖𝑛𝑖𝑚𝑢𝑚 𝑑𝑒𝑠𝑐𝑒𝑛𝑡 𝑟𝑎𝑡𝑒 𝑓𝑟𝑜𝑚 𝑎𝑝𝑜𝑔𝑒𝑒 𝑡𝑜 𝑚𝑎𝑖𝑛 𝑑𝑒𝑝𝑙𝑜𝑦𝑚𝑒𝑛𝑡 = 65.8924𝑓𝑡/𝑠
Tarleton Aeronautical Team 60
2012 - 2013 USLI Preliminary Design Review
Thus the minimum descent rate from apogee to main deployment in order to keep
total drift below 2500 feet is 65.8924 feet per second. With this figure it is possible
to calculate a maximum drogue parachute size that causes the descent rate to be
less than 65.8924 feet per second. The following calculates the maximum
diameter of the drogue parachute:
𝐴𝑠𝑠𝑢𝑚𝑒 𝑡ℎ𝑒 𝑣𝑒ℎ𝑖𝑐𝑙𝑒 𝑖𝑠 𝑑𝑒𝑠𝑐𝑒𝑛𝑑𝑖𝑛𝑔 𝑎𝑡 𝑎 𝑐𝑜𝑛𝑠𝑡𝑎𝑛𝑡 𝑠𝑝𝑒𝑒𝑑 (𝑠𝑡𝑒𝑎𝑑𝑦 𝑠𝑡𝑎𝑡𝑒).
𝐴𝑠𝑠𝑢𝑚𝑒 𝑡ℎ𝑒 𝑣𝑒ℎ𝑖𝑐𝑙𝑒 𝑠𝑖𝑚𝑝𝑙𝑦 𝑚𝑜𝑣𝑒𝑠 𝑑𝑜𝑤𝑛𝑤𝑎𝑟𝑑 (𝑐𝑜𝑛𝑠𝑡𝑟𝑎𝑖𝑛𝑒𝑑 𝑡𝑜 𝑧 − 𝑎𝑥𝑖𝑠).
𝐷 = 𝑑𝑟𝑎𝑔;
𝑊 = 𝑤𝑒𝑖𝑔ℎ𝑡;
∑ -𝐹𝑧 = 𝐷 − 𝑊 = 𝑚𝑔 = 0
𝑝 = 𝑑𝑒𝑛𝑠𝑖𝑡𝑦 𝑜𝑓 𝑎𝑖𝑟 =
0.075𝑙𝑏
;
𝑓𝑡 3
𝑆 = 𝑠𝑢𝑟𝑓𝑎𝑐𝑒 𝑎𝑟𝑒𝑎 𝑜𝑓 𝑡ℎ𝑒 𝑝𝑎𝑟𝑎𝑐ℎ𝑢𝑡𝑒
𝐶𝑑 = 𝑐𝑜𝑒𝑓𝑓𝑖𝑐𝑖𝑒𝑛𝑡 𝑜𝑓 𝑑𝑟𝑎𝑔 ≈ 2
𝐷=𝑊=
1 2
𝑝𝑣 𝑆𝐶𝑑
2
𝑆𝑜𝑙𝑣𝑒 𝑓𝑜𝑟 𝑠𝑢𝑟𝑓𝑎𝑐𝑒 𝑎𝑟𝑒𝑎 𝑜𝑓 𝑡ℎ𝑒 𝑝𝑎𝑟𝑎𝑐ℎ𝑢𝑡𝑒.
𝑆=
𝑆=
2𝑊
𝑝𝑣 2 𝐶𝑑
2 ∙ (29.62𝑙𝑏𝑠)
32.2𝑙𝑏
∙
3
2
(0.075𝑙𝑏/𝑓𝑡 ) ∙ (65.8924𝑓𝑡/𝑠) ∙ 2 1𝑙𝑏𝑓𝑠 2 /𝑓𝑡
𝑆 = 2.9289𝑓𝑡 2
𝐴 = 𝑎𝑟𝑒𝑎 𝑜𝑓 𝑎 𝑐𝑖𝑟𝑐𝑙𝑒
𝐴 = 𝜋𝑟 2
𝑆𝑜𝑙𝑣𝑒 𝑓𝑜𝑟 𝑑𝑖𝑎𝑚𝑒𝑡𝑒𝑟.
𝑑𝑖𝑎𝑚𝑒𝑡𝑒𝑟 = 2 ∙ √
=2∙√
3.079𝑓𝑡 2
𝜋
𝐴
𝜋
Tarleton Aeronautical Team 61
2012 - 2013 USLI Preliminary Design Review
𝑚𝑎𝑥𝑖𝑚𝑢𝑚 𝑑𝑖𝑎𝑚𝑒𝑡𝑒𝑟 𝑜𝑓 𝑡ℎ𝑒 𝑑𝑟𝑜𝑔𝑢𝑒 𝑝𝑎𝑟𝑎𝑐ℎ𝑢𝑡𝑒 = 1.9311𝑓𝑡
Based on the above calculations, commercial availability, and an overestimate of
5400 feet for apogee, the selected drogue parachute diameter is 24 inches.
Parachute
Drogue
Main
Diameter
2 ft
10 ft
Table 22 Parachute Diameters
The parachutes are 1.1 ounce silicon-coated nylon to ensure durability and
strength. Basing information on the resilient construction of the parachutes, flight
simulations from OpenRocket (see table 26), and commercially available size
limits, the diameter of the drogue is 2 feet, while that of the main 10 feet.
Recovery Component Itemization
Parachute Selection
Of the commercially available parachutes in consideration, those which made it
into the trade and selection process have the least number of shroud lines to limit
the probability of entanglement upon ejection, as well as reinforced seams with
nylon webbing to reduce the probability of shroud line disconnection as a result of
ejection. The selections of the round nylon parachutes for each the main and
drogue parachutes is in Table 23 and Table 24.
Item
Shroud
Lines
Diameter
Unit
Cost
TFR Par-120
16
10ft
$89.95
The Rocketman
STANDARD LOW-POROSITY 1.1
RIPSTOP PARACHUTE
4
10ft
Fruity Chutes
Custom Parachute - 50lb @ 20fps
16
10ft
Manufacturer
Top Flight
Recovery
Table 23 - Main Parachute Trade and Selection
$110.0
0
$240.0
0
Tarleton Aeronautical Team 62
2012 - 2013 USLI Preliminary Design Review
Manufacturer
Top Flight
The
Rocketman
Apogee
Rockets
Item
PAR-STD-24
STANDARD LOW-POROSITY 1.1
RIPSTOP
Shroud
Lines
6
Diamete
r
2ft
Unit
Cost
$9.95
4
2ft
$25.00
6
2ft
$11.44
Nylon Parachute #29218
Table 24 - Drogue Parachute Trade and Selection
Parachute Protection
Each parachute is quick-linked to the bridle of a 2,000 degrees Fahrenheit-plusheat-resistant Nomex deployment bag. Cellulose (“dog barf”) insulation and sheet
insulation wadding were also viable candidates for their wide availability and low
cost. Through testing, it is determinable whether dog barf is usable in conjunction
with the deployment bag system.
However, hand-packed insulation such as dog barf or sheet wadding could easily
shift during ascent, potentially exposing some portion of the parachute, shroud
lines, or shock cord. This could result in a fire hazard and entanglement. The
deployment bag keeps the parachutes, shroud lines, and shock cord away from
the heat of the firing ejection charge reliably throughout ascent and allows for
ejection in an orderly fashion. Table 25 demonstrates this selection.
Component
Flameproof Main Parachute
Deployment Bag
Recovery Wadding for Main
Compartment
Dog Barf for Main
Compartment
Flame-Proof Drogue
Parachute Deployment Bag
Recovery Wadding for
Drogue Compartment
Dog Barf for Drogue
Compartment
Distributor
Rocketman
Enterprises
Item
HobbyLinc
Rockets "R"
Us
Rocketman
Enterprises
EST302274
"Dog Barf" Recovery
Wadding - 16oz
HobbyLinc
Rockets "R"
Us
EST302274
"Dog Barf" Recovery
Wadding - 16oz
DB8
DB2
Table 25 - Parachute Protection Materials
Unit Dimensions
3.25in. X 3.25in. X
6.5in. (custom)
Unit
Cost
$40.00
75 sheets
$3.69
16oz
3.25in. X 3.25in. X
18.5in. (custom)
$7.00
$25.00
75 sheets
$3.69
16oz
$7.00
Tarleton Aeronautical Team 63
2012 - 2013 USLI Preliminary Design Review
It is important to know that these parachutes have a probability of entanglement,
as the main parachute has 16 shroud lines; the drogue has six. To counteract this
issue, standard high-power model rocketry 900 pound-working-strength swivels
are on each end of each shock harness, connecting each shock harness to both
the parachute and the eyebolt. Each parachute is neatly folded into its deployment
bag with all shroud lines and shock cord. This keeps each parachute inside long
enough to maintain separation from the fins and to ensure that ejection takes place
in an orderly fashion.
Shock Chord
In order to enable deceleration of the two tethered sections on descent, 0.5-inch
tubular Kevlar shock harnesses are necessary. These implementations are for
strength and flame-proof construction. Additionally, these shock harnesses include
pre-sewn Nomex loops for a safe and secure connection to the parachutes by
delta-shaped quick links.
The main harness is to be approximately two body lengths plus fifteen percent to
keep the main parachute away from the body. The drogue harness is to be
approximately three body lengths plus fifteen percent to pull the drogue parachute
up and away from the vehicle, while minimizing the risk of denting or zippering.
The body length is approximately nine feet, including the nose cone. Thus, based
on commercial availability, the length of the main shock harness is to be 20 feet,
while that for the drogue is to be 25 feet. Table 26 illustrates this selection.
Component
Main Shock Cord
Drogue Shock
Cord
Unit
Dimensions
0.5in. X 25ft
Unit
Cost
$37.99
1in. X 20ft
$30.00
Tubular Kevlar
0.5in. X 20ft
$31.49
Bulk Tubular Nylon
Webbing
1in. X 15ft
$25.00
Distributor
Item
Giant Leap Rocketry
Rocketman
Enterprises
Tubular Kevlar
Bulk Tubular Nylon
Webbing
Giant Leap Rocketry
Rocketman
Enterprises
Table 26 - Shock Chord Trade and Selection
Stainless steel delta-shaped quick links, each with a working load of 1,000 pounds,
are a in the design. These secure each parachute to its shock harness and each
shock harness to a U-bolt at the bulkhead of each altimeter compartment to
ensure secure tethering throughout the duration of the flight.
Tarleton Aeronautical Team 64
2012 - 2013 USLI Preliminary Design Review
Shear Pins
In order to ensure that separation in the rocket body only occurs upon ejection, the
design makes use of removable threaded nylon shear pins. These are inserted
through holes on either side of the couplings between the nose cone shoulder and
main parachute compartment, as well as between the payload and drogue
parachute compartments. When the ejection charge fires, the force of the coupler
sliding past will snap the shear pins. However, other stresses under 25 pound
force such as those caused by shifting mass, drag, or ejection from another
compartment should not be strong enough to cause separation.
The itemized components of the recovery section are in the following Table 27.
This includes the total cost and mass (where available) of each item. The total cost
reflects one fully-assembled rocket.
Table 26 – Launch-Day Recovery System Budget
Proposed
Selection
Distributor
Item Number
Unit
Dimensions
Unit
Cost
Qty.
Main
Altimeters
Featherweight
Altimeters
Raven3
1.8in. X 0.8in.
X 0.55in. X
0.34oz
$155.00
2
$310.00
Backup
Altimeters
PerfectFlite
StratoLogger
2.75in. X
0.9in. 0.55in.,
0.45oz
$79.95
2
$159.90
Coast Rocketry
Daveyfire
N28BR
1ft long
$2.95
4
$11.80
FFFFg Black
Powder
Goex
Goex 4F Black
Powder
1lb
$15.75
1
$15.75
Black Powder
Ejection
Charge
Holders
Aerocon
Systems
BPSmall
15 X 0.067oz
$3.00
1
$3.00
Commonwealth
Rocketry
SWLDK80
1.5in.
$1.99
2
$3.98
Main Shock
Cord
Giant Leap
Rocketry
Tubular Kevlar
0.5in. X 25ft
$37.99
1
$37.99
Drogue Shock
Cord
Giant Leap
Rocketry
Tubular Kevlar
0.5in. X 20ft
$31.49
1
$31.49
Main
Parachute
Top Flight
Recovery
TFR Par-120
10ft Diameter
$89.95
1
$89.95
Rocketman
Enterprises
DB8
3.25in. X
3.25in. X
$40.00
1
$40.00
Electric
Matches
Swivels
Flameproof
Main
Total Cost
Tarleton Aeronautical Team 65
2012 - 2013 USLI Preliminary Design Review
Parachute
Deployment
Bag
Drogue
Parachute
6.5in.
(custom)
Top Flight
PAR-STD-24
2ft Diameter
$9.95
1
$9.95
$25.00
1
$25.00
Flame-Proof
Drogue
Parachute
Deployment
Bag
Rocketman
Enterprises
DB2
3.25in. X
3.25in. X
18.5in.
(custom)
U-Bolts
Sunward
Aerospace
U-Bolt
Assembly 0.25in.
(compact)
0.212in. X
1.5in. X 2.125
X 1.77oz
$4.29
2
$8.58
Quick Links
Commonwealth
Rocketry
0.25in.
Stainless Steel
Delta Quick
Link
2.375in. X
0.375in. X
1.25in.
$2.99
4
$11.96
Shear Pins
Missile Works
2-56 Nylon
Shear-Pin (10
pack)
0.08in. X
0.5in. X
0.002oz X 10
$1.00
1
$1.00
Arming
Switches
Featherweight
Altimeters
Featherweight
Magnetic
Switch
0.55in. X
0.75in.
$25.00
2
$50.00
Hand-held
Rotary Fan
Anemometer
Weather Shack
SpeedTech
WM-300
5.5in. X 1.7in.
X .75in.
$154.95
1
$154.95
GPS System
Big Red Bee
BeeLine GPSPackage Deal
5.5in. X
1.75in. X
0.75in.
$289.00
2
$578.00
LessEMF
Pure Copper
Polyester
Taffeta Fabric
12in. X 42.5in.
X 0.003 in.
$10.95
2
$21.90
Radio
Frequency
Shielding
Material
Total
$1,565.20
Table 27 Launch Day Recovery System Budget
Tarleton Aeronautical Team 66
2012 - 2013 USLI Preliminary Design Review
Mission Performance Predictions
Mission Performance Criteria
The rocket vehicle delivers the SMD payload to 5,280 feet above ground level. It
carries a barometric altimeter for official scoring, and must remain subsonic during
the entire flight, from launch until landing. The vehicle must be recoverable and
reusable on the day of the official launch.
Simulations
Flight Simulations
Performances of flight simulations in OpenRocket help gather data concerning
expectations for the vehicle. All relevant parameters are input into this simulation
to acquire vehicle performance data including component weights. The mass of
the rocket is slightly over-estimated to account for weight gain as the design
progresses. The simulations account for launch conditions using a yearly average
for the date and location of the official launch site. Launch conditions include 9±.9
mph average wind speed, where ±.9 mph was found using 10% turbulence
intensity. A 75° F temperature and 1241 mbar pressure is in place for the
simulation as well. All weather data accumulation is from the National Oceanic and
Atmospheric Administration (NOAA). Latitude and longitude locations of 34.9° N
and 86.5°W are applicable for the project.
Tarleton Aeronautical Team 67
2012 - 2013 USLI Preliminary Design Review
Altitude
With these launch conditions, the vehicle achieves a height of 5350±5 feet. The
following graphs illustrate the altitude throughout the simulated flight. After 2.11
seconds of motor burn, the estimate for altitude of the vehicle is 760 feet. The
vehicle continues to climb to apogee at 5,245 feet where the initial separation
event and drogue parachute deployment occur. This precedes the main parachute
deployment at 500 feet.
Figure 20 Altitude Simulation
Tarleton Aeronautical Team 68
2012 - 2013 USLI Preliminary Design Review
Velocity
Under guidance from the team mentor, Pat Gordzelik, an exit velocity over 60 feet
per second is the goal to ensure adequate stability when leaving the launch rail.
Additionally, the intent is for the vehicle to remain subsonic to meet the
requirements set forth in the Statement of Work (SOW). The following graph
shows the entire simulation of the velocity from launch until landing. After 0.24
seconds of motor burn, the velocity of the vehicle when departing from the launch
rail should be 76 feet per second; this exceeds the recommended 60 feet per
second. After a total of 2.11 seconds when the motor propellant depletes, a
maximum vertical velocity of 833 feet per second is the result; this demonstrates
that the vehicle remains subsonic throughout the flight. Apogee should occur after
17.68 seconds into the flight, with a theoretical altitude of 5,345 feet. Drogue
parachute deployment occurs at apogee and the vehicle begins to descend at a
velocity of approximately 75 feet per second until the main parachute deployment
occurs at 500 feet. At this time, the velocity decreases to 17.7 feet per second until
landing.
Figure 21 Velocity Simulation
Tarleton Aeronautical Team 69
2012 - 2013 USLI Preliminary Design Review
Acceleration
Throughout the flight, multiple stages of acceleration occur. Two graphs illustrate
the acceleration of the vehicle throughout the flight. The first stage represents the
acceleration from ignition to burnout of the motor. The maximum acceleration of
the motor is approximately 400 feet per second per second. This occurs 0.789
seconds into the flight. At approximately 2.11 seconds into the flight, just prior to
motor burnout, the acceleration decreases.
Figure 22 Acceleration Before Burn Out Simulation
Tarleton Aeronautical Team 70
2012 - 2013 USLI Preliminary Design Review
The second graph represents the simulated acceleration from motor burn out to
landing. At apogee, the acceleration is -31.28 feet per squared second. At an
altitude of 500 feet, the main parachute deploys, and the vehicle momentarily
accelerates to 553 feet per squared second for less than .1 second. The
acceleration then falls into an oscillation, ranging from -.15 to .15 feet per squared
second for the remainder of the descent to ground level.
Figure 23 Acceleration After Burn Out
Tarleton Aeronautical Team 71
2012 - 2013 USLI Preliminary Design Review
Figure 24 is the simulated thrust curve for the Cesaroni L1720-WT motor.
Figure 24 L1720 Thrust Curve
Stability
Center of Pressure/Gravity
The stability margin should be at least two calibers of body diameter. The center of
gravity is 65.667 inches from the nose, while the center of pressure is at 79.738
inches from the nose. This provides 14.071 inches of stability margin or 2.55
calibers of body width.
Figure 25 Stability: Center of Pressure/Gravity
Tarleton Aeronautical Team 72
2012 - 2013 USLI Preliminary Design Review
Kinetic Energy
The weight of the rocket upon landing is 29.62 pounds. The weight of the nose
cone section is approximately 2.252 pounds, while that of the upper body airframe
and payload is approximately 16.766 pounds, and the booster section is
approximately 10.602 pounds, excluding propellant.
The following calculates the kinetic energy, first for the entire rocket, then for each
individual section. First it is necessary to calculate the descent velocity upon
landing.
𝑆 = 𝐴𝑟𝑒𝑎 𝑜𝑓 𝑚𝑎𝑖𝑛 𝑝𝑎𝑟𝑎𝑐ℎ𝑢𝑡𝑒 = 78.54𝑓𝑡 2 ;
𝑆=
𝑊 = 𝑤𝑖𝑒𝑔ℎ𝑡 = 29.62𝑙𝑏𝑠
2𝑊
𝑝𝑣 2 𝐶𝑑
𝑆𝑜𝑙𝑣𝑒 𝑓𝑜𝑟 𝑑𝑒𝑠𝑐𝑒𝑛𝑡 𝑟𝑎𝑡𝑒.
𝑣=√
= √
2𝑊
𝑝𝐶𝑑 𝑆
2 ∙ 29.62𝑙𝑏
32.2𝑙𝑏
∙
3
2
(0.075𝑙𝑏/𝑓𝑡 ) ∙ 2 ∙ 78.5398𝑓𝑡 1𝑙𝑏𝑓𝑠 2 /𝑓𝑡
𝑑𝑒𝑠𝑐𝑒𝑛𝑡 𝑟𝑎𝑡𝑒 = 12.7246𝑓𝑡/𝑠
Using 12.7246 feet per second for the landing velocity it is possible to calculate the
kinetic energy for the entire vehicle and each individual section, with relation to
their mass. The following calculates the kinetic energy of the entire vehicle.
𝑣 = 𝑑𝑒𝑠𝑐𝑒𝑛𝑡 𝑟𝑎𝑡𝑒 = 12.7246
𝑤𝑒𝑖𝑔ℎ𝑡 = 𝑚𝑔 = 29.62𝑙𝑏𝑠
𝑔 = 𝑔𝑟𝑎𝑣𝑖𝑡𝑦 = 32.2𝑓𝑡/𝑠 2
𝐾𝑖𝑛𝑒𝑡𝑖𝑐 𝐸𝑛𝑒𝑟𝑔𝑦 = 𝐾𝐸 =
1
𝑚𝑣 2
2
Tarleton Aeronautical Team 73
2012 - 2013 USLI Preliminary Design Review
𝐾𝐸 =
29.62𝑙𝑏𝑠 ∙ (12.7246)2
2 ∙ (32.2𝑓𝑡/𝑠 2 )
𝐾𝑖𝑛𝑒𝑡𝑖𝑐 𝐸𝑛𝑒𝑟𝑔𝑦 = 74.471𝑓𝑡 ∙ 𝑙𝑏𝑓
Therefore the kinetic energy for the entire vehicle as a whole is 74.471 foot pound
force. This leads to the assumption that if the kinetic energy of the entire vehicle is
74.471 foot pound force, then the kinetic energy for each individual section must
be less than 74.471 foot pound force. It is possible to prove this assumption by
calculating the kinetic energy of each individual section. The following calculates
the kinetic energy of the nose cone, which has a weight of 2.252 pounds.
𝑣 = 𝑑𝑒𝑠𝑐𝑒𝑛𝑡 𝑟𝑎𝑡𝑒 = 12.7246 𝑓𝑡/𝑠
𝑛𝑜𝑠𝑒 𝑐𝑜𝑛𝑒 𝑤𝑒𝑖𝑔ℎ𝑡 = 𝑚𝑔 = 2.252𝑙𝑏𝑠
𝑔 = 𝑔𝑟𝑎𝑣𝑖𝑡𝑦 = 32.2𝑓𝑡/𝑠 2
𝐾𝑖𝑛𝑒𝑡𝑖𝑐 𝐸𝑛𝑒𝑟𝑔𝑦 = 𝐾𝐸 =
𝐾𝐸 =
1
𝑚𝑣 2
2
2.252𝑙𝑏𝑠 ∙ (12.7246)2
2 ∙ (32.2𝑓𝑡/𝑠 2 )
𝐾𝑖𝑛𝑒𝑡𝑖𝑐 𝐸𝑛𝑒𝑟𝑔𝑦 = 5.662𝑓𝑡 ∙ 𝑙𝑏𝑓
The following math calculates the kinetic energy of the connected upper body air
frame and payload. This section has a weight of 16.766 pounds.
𝑣 = 𝑑𝑒𝑠𝑐𝑒𝑛𝑡 𝑟𝑎𝑡𝑒 = 12.7246 𝑓𝑡/𝑠
𝑢𝑝𝑝𝑒𝑟 𝑏𝑜𝑑𝑦 𝑎𝑖𝑟 𝑓𝑟𝑎𝑚𝑒 𝑎𝑛𝑑 𝑝𝑎𝑦𝑙𝑜𝑎𝑑 𝑤𝑒𝑖𝑔ℎ𝑡 = 𝑚𝑔 = 16.766𝑙𝑏𝑠
𝑔 = 𝑔𝑟𝑎𝑣𝑖𝑡𝑦 = 32.2𝑓𝑡/𝑠 2
𝐾𝑖𝑛𝑒𝑡𝑖𝑐 𝐸𝑛𝑒𝑟𝑔𝑦 = 𝐾𝐸 =
1
𝑚𝑣 2
2
16.766𝑙𝑏𝑠 ∙ (12.7246)2
𝐾𝐸 =
2 ∙ (32.2𝑓𝑡/𝑠 2 )
𝐾𝑖𝑛𝑒𝑡𝑖𝑐 𝐸𝑛𝑒𝑟𝑔𝑦 = 42.1533𝑓𝑡 ∙ 𝑙𝑏𝑓
Tarleton Aeronautical Team 74
2012 - 2013 USLI Preliminary Design Review
The following calculates the kinetic energy of the booster section. This section has
a weight of 10.602 pounds.
𝑣 = 𝑑𝑒𝑠𝑐𝑒𝑛𝑡 𝑟𝑎𝑡𝑒 = 12.7246 𝑓𝑡/𝑠
𝑤𝑒𝑖𝑔ℎ𝑡 = 𝑚𝑔 = 10.602𝑙𝑏𝑠
𝑔 = 𝑔𝑟𝑎𝑣𝑖𝑡𝑦 = 32.2𝑓𝑡/𝑠 2
𝐾𝑖𝑛𝑒𝑡𝑖𝑐 𝐸𝑛𝑒𝑟𝑔𝑦 = 𝐾𝐸 =
𝐾𝐸 =
1
𝑚𝑣 2
2
10.602𝑙𝑏𝑠 ∙ (12.7246)2
2 ∙ (32.2𝑓𝑡/𝑠 2 )
𝐾𝑖𝑛𝑒𝑡𝑖𝑐 𝐸𝑛𝑒𝑟𝑔𝑦 = 26.6557𝑓𝑡 ∙ 𝑙𝑏𝑓
Table 28 summarizes the kinetic energy for each individual section and the vehicle
as a whole.
Section
Weight
Kinetic Energy
Nose Cone
2.252 lbs
5.662 ft∙lbf
Upperbody Airframe and Payload
16.766 lbs
42.1533 ft∙lbf
Booster
10.602 lbs
26.6557 ft∙lbf
Total
29.62 lbs
74.471 ft∙lbf
Table 28 Kinetic Energy Summarization
Drift Calculations
Wind causes two effects during flight, drift after parachute deployment and
weather cocking during vertical ascent. The parachute selection calculations do
not take in to account weather cocking but rather calculate the drift after parachute
deployment. The following formula is the basis for the drift calculations. The
formula takes in to account the two different descent rates, one for when the
drogue parachute is deployed and one for when the main parachute is deployed.
The apogee altitude is approximated to 5280 feet.
∆𝐴𝑑 = 𝑐ℎ𝑎𝑛𝑔𝑒 𝑖𝑛 𝑎𝑙𝑡𝑖𝑡𝑢𝑑𝑒 𝑑𝑢𝑟𝑖𝑛𝑔 𝑑𝑟𝑜𝑔𝑢𝑒 𝑑𝑒𝑝𝑙𝑜𝑦𝑚𝑒𝑛𝑡 = 4780𝑓𝑡
∆𝐴𝑚 = 𝑐ℎ𝑎𝑛𝑔𝑒 𝑖𝑛 𝑎𝑙𝑡𝑖𝑡𝑢𝑑𝑒 𝑑𝑢𝑟𝑖𝑛𝑔 𝑡ℎ𝑒 𝑚𝑎𝑖𝑛 𝑑𝑒𝑝𝑙𝑜𝑦𝑚𝑒𝑛𝑡 = 500𝑓𝑡
𝑣𝑑 = 𝑑𝑟𝑜𝑔𝑢𝑒 𝑑𝑒𝑠𝑐𝑒𝑛𝑡 𝑟𝑎𝑡𝑒 = 63.6231;
𝑣𝑚 = 𝑚𝑎𝑖𝑛 𝑑𝑒𝑠𝑐𝑒𝑛𝑡 𝑟𝑎𝑐𝑒 = 12.7246𝑓𝑡/𝑠
𝑣ℎ = ℎ𝑜𝑟𝑖𝑧𝑜𝑛𝑡𝑎𝑙 𝑤𝑖𝑛𝑑 𝑣𝑒𝑙𝑜𝑐𝑖𝑡𝑦
𝐷𝑟𝑖𝑓𝑡 =
∆𝐴𝑑
∆𝐴𝑚
∙ 𝑣ℎ +
∙ 𝑣ℎ
𝑣𝑑
𝑣𝑚
Tarleton Aeronautical Team 75
2012 - 2013 USLI Preliminary Design Review
The following calculates the theoretical landing radius of the vehicle for zero miles
per hour. These results do not take into account weather cocking.
𝐷𝑟𝑖𝑓𝑡 =
4780𝑓𝑡
500𝑓𝑡
∙ 0𝑓𝑡/𝑠 +
∙ 0𝑓𝑡/𝑠
62.6231𝑓𝑡/𝑠
12.7246𝑓𝑡/𝑠
𝐷𝑟𝑖𝑓𝑡 𝑤𝑖𝑡ℎ 0𝑚𝑝ℎ 𝑤𝑖𝑛𝑑 𝑠𝑝𝑒𝑒𝑑 = 0𝑓𝑡
For five miles per hour (7.33 feet per second):
𝐷𝑟𝑖𝑓𝑡 =
4780𝑓𝑡
500𝑓𝑡
∙ 7.33𝑓𝑡/𝑠 +
∙ 7.33𝑓𝑡/𝑠
62.6231𝑓𝑡/𝑠
12.7246𝑓𝑡/𝑠
𝐷𝑟𝑖𝑓𝑡 𝑤𝑖𝑡ℎ 5𝑚𝑝ℎ 𝑤𝑖𝑛𝑑 𝑠𝑝𝑒𝑒𝑑 = 847.521𝑓𝑡
For ten miles per hour (14.67 feet per second):
𝐷𝑟𝑖𝑓𝑡 =
4780𝑓𝑡
500𝑓𝑡
∙ 14.67𝑓𝑡/𝑠 +
∙ 14.67𝑓𝑡/𝑠
62.6231𝑓𝑡/𝑠
12.7246𝑓𝑡/𝑠
𝐷𝑟𝑖𝑓𝑡 𝑤𝑖𝑡ℎ 10𝑚𝑝ℎ 𝑤𝑖𝑛𝑑 𝑠𝑝𝑒𝑒𝑑 = 1696.199𝑓𝑡
For fifteen miles per hour (22 feet per second):
𝐷𝑟𝑖𝑓𝑡 =
4780𝑓𝑡
500𝑓𝑡
∙ 22𝑓𝑡/𝑠 +
∙ 22𝑓𝑡/𝑠
62.6231𝑓𝑡/𝑠
12.7246𝑓𝑡/𝑠
𝐷𝑟𝑖𝑓𝑡 𝑤𝑖𝑡ℎ 15𝑚𝑝ℎ 𝑤𝑖𝑛𝑑 𝑠𝑝𝑒𝑒𝑑 = 2543.72𝑓𝑡
(𝑇𝐻𝐼𝑆 𝐷𝑂𝐸𝑆 𝑁𝑂𝑇 𝑇𝐴𝐾𝐸 𝐼𝑁𝑇𝑂 𝐴𝐶𝐶𝑂𝑈𝑁𝑇 𝑊𝐸𝐴𝑇𝐻𝐸𝑅 𝐶𝑂𝐶𝐾𝐼𝑁𝐺)
For twenty miles per hour wind (29.33 feet per second):
𝐷𝑟𝑖𝑓𝑡 =
4780𝑓𝑡
500𝑓𝑡
∙ 29.33𝑓𝑡/𝑠 +
∙ 29.33𝑓𝑡/𝑠
62.6231𝑓𝑡/𝑠
12.7246𝑓𝑡/𝑠
𝐷𝑟𝑖𝑓𝑡 𝑤𝑖𝑡ℎ 20𝑚𝑝ℎ 𝑤𝑖𝑛𝑑 𝑠𝑝𝑒𝑒𝑑 = 3391.24𝑓𝑡
The previous results did not take into account weather cocking. Weather cocking
occurs when a horizontal wind velocity rotates the vehicle to where it has a new
flight direction into the wind. Figure 26, found on the NASA website, demonstrates
the effects of weather cocking.
Tarleton Aeronautical Team 76
2012 - 2013 USLI Preliminary Design Review
Figure 26 Weather Cocking
Weather cocking causes the vehicle to fly into the wind. Then after apogee,
horizontal wind velocity causes the vehicle to drift with the wind. This effect greatly
decreases the landing radius when compared to theoretical calculations. Open
Rocket simulations calculate weather cocking in flight trajectory.
Figure 27 is a plot of the vehicle’s distance from the launch pad in relation to
altitude. The simulation assumes a 15 mile per hour wind.
Tarleton Aeronautical Team 77
2012 - 2013 USLI Preliminary Design Review
Figure 27 Weather Cocking Simulation
Table 29 summarizes the simulated landing radius of the vehicle for five different
wind speeds. These calculations take in to account both weather cocking and
parachute drift.
Wind Speed
Landing Radius
0 mph
7 ft
5 mph (7.33 ft/s)
275 ft
10 mph (14.67 ft/s)
550 ft
15 mph (22 ft/s)
900 ft
20 mph (29.33 ft/s)
1300 ft
Table 29 Landing Radius
Tarleton Aeronautical Team 78
2012 - 2013 USLI Preliminary Design Review
Interfaces and Integration
Payload Integration Plan
The payload is designed to integrate into the payload housing structure of the
launch vehicle in a simple and easy fashion. The payload is constructed on a
payload framework which consists of two bulkheads connected to each other by
two aluminum rails. Because the payload housing structure is part of the vehicle,
the payload framework uses dimensions to ensure compatibility between the
payload and the vehicle.
Figure 28 Payload Pre-Integration
The payload is built onto the payload framework, and preparing the payload for
flight is easily done prior to payload framework integration with the vehicle. At this
time the framework is inserted from the lower portion of the payload housing
structure. Once the payload framework is in place, it is secured with the use of two
screws; from the exterior of the payload housing structure and into the bulkhead.
At this point the physical integration of payload is complete.
Tarleton Aeronautical Team 79
2012 - 2013 USLI Preliminary Design Review
Internal Vehicle Interfaces
Structure
Structure(heading lv 4) (upper body structure=upper body airframe; lower body
structure= booster section)
The flight vehicle consists of four main sections; the nose cone, the upper body
airframe, the payload housing structure, and the booster section. These sections
interface with couplers, rivets, screws, and shear pins. Two identical 11 inch long
couplers are modified to have a larger outer diameter on one side. This
modification allows compatibility of the inner diameter of the acrylic payload
housing structure. Each coupler interfaces the payload housing structure with the
fiberglass upper body airframe and booster section respectively. Figure 29 shows
the coupler specifications.
Figure 29 Coupler Specifications
Tarleton Aeronautical Team 80
2012 - 2013 USLI Preliminary Design Review
The nose cone shoulder is used to couple the nose cone with the upper body
airframe. This coupling is secured with shear pins, as the nose cone must
separate from the upper body airframe to deploy the main parachute. A Kevlar
shock cord is used to tether the nose cone and the upper body airframe. The
Kevlar tether is attached between two eyebolts.
The upper body airframe and the payload housing structure are coupled with a
fiberglass coupler and four aluminum rivets; two rivets through the upper body
airframe, and two additional rivets through the payload housing structure. The
rivets attaching the payload housing structure with the coupler will remain
unaltered for the life of the vehicle; whereas, the rivets connecting the upper body
airframe to the coupler may be drilled out and reinstalled to gain access to the
components in the lower portion of the upper body airframe, namely the GPS bay
and main parachute altimeter bay.
The payload housing structure, from the bottom end, couples with the booster
section using a fiberglass coupler, four screws, and two shear pins. The screws
will attach the bottom portion of the payload housing structure to the fiberglass
coupler and the bottom bulkhead of the payload framework. The booster section
and the coupler attach as a friction fit, but also utilize two shear pins. Additionally
there is an eyebolt anchored at the bottom of the payload framework bulkhead,
which secure to the payload housing structure. A Kevlar shock chord tethers the
payload housing structure eyebolt to another eyebolt in the booster section.
For the booster section, fins mount through slits in the vehicle body and epoxy in
place. The motor mount tube attaches with centering rings to the inner surface of
the booster section and epoxy in place.
Launch Vehicle to Ground Interfaces
The recovery system altimeter arming switches are a mechanical interface with the
ground. Physically removable magnetic pins activate the arming switches upon
removal. The pins are not removed until the vehicle is on the launch pad and ready
for flight. Upon removal, the altimeters audibly respond with arming confirmation.
The drogue deployment altimeter arming switch is located two feet above the base
of the vehicle and the main deployment altimeter is located five feet eight inches
above the base of the launch vehicle, which does not exceed six feet
(Requirement 2.9). The BeeLine GPS and the XBee XSC S3B radio, payload
section, are wireless transmission interfaces between the ground and the vehicle.
Neither wireless device receive telemetry, but both transmit telemetry to the
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2012 - 2013 USLI Preliminary Design Review
ground station. The BeeLine GPS transmits readings of latitude and longitude for
recovery purposes and the XBee XSC s3B transmits atmospheric data. Table 30
summarizes the interfaces between the vehicle and the ground.
Ground – Vehicle Interface
Device
Mechanical
Arming Switches
Wireless Transmission
BeeLine GPS, Xbee XSC S3B
Table 30 Ground- Vehicle Interface
Vehicle to Ground Launch System Interfaces
Rail buttons are attached to the vehicle to provide an interface between the vehicle
and the launch rail. The selected rail buttons function with a 1010 launch rail. The
bottom rail button is located two inches from the base of the vehicle and the upper
rail button is located fifteen inches from the base of the vehicle. The motor ignition
system interfaces the vehicle and the ground launch system. The motor ignites by
a standard twelve volt firing system. The igniter is provided with the Cesaroni
L1720-WT-P motor.
Launch Operation Procedures
Appendix B contains checklists for the final assembly and launch procedures. The
checklists provide a detailed description of the steps to ensure a successful flight.
The checklists will be strictly followed and updated as needed.
Safety and Environment (Vehicle)
The Safety Officer
The team safety officer, Blake, is level one certified with NAR. The responsibility of
the safety officer is to design and implement safety plans that ensure all accidents
are evaded. All hazards to people, the project, and the mission are determined so
that mitigations can be enacted.
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The systematic identification of risks, failure modes, and personnel hazards allows
for the team to discover where single points of failure could occur throughout the
course of the project. The identification of single point failures allows for proactive
design changes to be made to counter these failures.
Failure Modes
A failure mode is the way in which a system could fail, causing an undesirable
effect on some aspect of the project. The safety plan ensures development and
implementation of mitigations for each failure mode.
Rocket Design Failure Modes
Table 31 provides a summary of potential failure modes that could occur during
the design of the vehicle.
Failure
Vehicle Unstable
Acrylics Does Not
Withstand Forces
Throughout Flight
Fiberglass Does
not Withstand
Forces
Throughout Flight
Connection
Between Acrylic
Payload Housing
and Upper
Fiberglass Body
Tube Becomes
Detached
Fins Cause To
Much Drag
Thrust To Weight
Ratio is Less Than
5:1
Effect
Proposed Mitigation
Completed Mitigation
Unpredictable
Flight Path
Simulations
Completed
Flight Testing
Proposed (12/1/2012)
Vehicle Not
Reusable
Tensile Strength and
Flight Testing
Proposed (12/1/2012)
Vehicle Not
Reusable
Tensile Strength and
Flight Testing
Proposed (12/1/2012)
Unpredictable
Flight Path,
Damage to
Vehicle Body
Research/ Design
Completed
Testing
Proposed (12/1/2012)
Simulations
Completed
Flight Testing
Proposed (12/1/2012)
Simulations/
Calculations
Completed
Expected
Apogee Height
Not Obtained
Unpredictable
Flight Path
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Couplers Too
Long or Too Short
Early or No
Separation
Research/
Simulations/
Calculations
Completed
Flight Testing
Proposed (12/1/2012)
Table 31 Potential Failure Modes for the Design of the Vehicle
Payload Integration Failure Modes
Table 32 provides a summary of potential failure modes that could occur during
payload integration.
Failure
Screw hole
stripped out
Rubber O-ring
tear
Circuitry
becomes hung
up
Effect
Proposed Mitigation
Inadequately
secured payload
Excessive
vibrations
Ensure the bulkhead is replaced
when required
Be sure to replace when
required
Electronic
malfunction
Be careful while inserting
payload
Completed Mitigation
Proposed (1/5/2012)
Proposed (1/5/2012)
Proposed (1/5/2012)
Table 32 Potential Failure Modes during Payload Integration
Launch Operations Failure Modes
Table 33 provides a summary of potential failure modes that could occur during
launch operations.
Subsystem
Structure
Potential Failure
Mode
Ejection charges
damage airframe/rocket
components
Potential Effects of
Failure
Proposed
Mitigations
Completed
Mitigations
Critical systems
become damaged
Calculations
Completed
Proper testing
Proposed
(1/5/2012)
Motor mount fails to
properly retain
motor
Damage to internal
systems
Rail button failure
Unpredictable flight
path
Insufficient
component
mounting
Airframe stress
failure
Structural testing
of the motor
mount
Ensure rail
buttons are
properly installed
and orientated
Proposed
(10/27/2012
Proposed
(1/5/2012)
Potential system
malfunction
Test mounting
integrity,
Proposed
(1/5/2012)
Loss of vehicle
functionality, potential
Structural testing
of airframe
Proposed
(1/5/2012)
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loss of vehicle
Fin Detachment
Aerodynamic
instability of vehicle
Potential Failure
Mode
Parachute shroud
line fails
Potential Effects of
Failure
Uncontrollable
descent
Electronic matches
do not fire
Eye bolt failure
Shock cord failure
Parachute
deployment does not
occur
Uncontrollable
descent
Untethered vehicle
components, violation
of requirements
Ensure that the
fins are properly
epoxied
Proposed
Mitigation
Premature black
power ignition
Premature parachute
ejection
Completed
Mitigations
Research
Completed
Verify parachute
rating
Proposed
(10/27/2012)
Redundant
altimeter system
Completed
Verify eye bolt
integrity
Proposed
(1/5/2012)
Properly fastened
shock cord
Proposed
(10/727/2012)
Verify rating
Recovery
Proposed
(12/1/2012)
Testing
Recovery
altimeter shielding
Use nomex cloth
and fire retardant
insulation
Proposed
(10/27/2012)
Proposed
(10/27/2012)
Completed
Recovery system
ignites
Failure of recovery
system, damage or
loss to vehicle
Main or drogue
parachute comes
untied from the
swivel
Uncontrolled descent
Secure swivels
along with quick
links.
Proposed
(1/5/2012)
Main or drogue
parachute shrouds
become entangled
Uncontrolled descent
rate
Ensure parachute
shroud lines are
attached to a
swivel
Proposed
(1/5/2012)
Failed Separation
PerfectFlites power
supply diminishes
Featherweight
Power supply
diminishes
PerfectFlite wired
connections
become damaged
from handling
Featherweight
wired connections
become damaged
Proposed
(10/27/2012)
Un-functional
recovery system,
ballistic descent
failure of deployment
of parachutes,
Mission failure
failure of deployment
of parachutes,
Mission failure
Adequate
separation testing
Proposed
(10/27/2012)
Use new batteries
before launch
Proposed
(1/5/2012)
Use new batteries
before launch
Proposed
(1/5/2012)
failure of deployment
of parachutes,
Mission failure
Use protected
electrical
components
Completed
failure of deployment
of parachutes,
Mission failure
Use protected
electrical
components
Completed
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from handling
Scoring Altimeter
failure
Potential Failure
Mode
Igniter does not
initiate the oxidation
process for the
propellant
Propulsion
We lose all points
associated with the
altitude portion of the
project
Potential Effects of
Failure
The rocket does not
launch
Propellant’s
oxidation process
does not
commence
The rocket does not
launch
A pressure build-up
occurs inside the
motor
Explosion
Redundant
Systems
Completed
Proposed
Mitigation
Completed
Mitigations
Inspect igniter for
concatenation
Proposed
(12/1/2012)
Always bring
additional igniters
for such an event
Use proper
igniter, sue
appropriate
conditions when
storing propellant
Inspect the motor
Table 33 Potential Failure Modes during Launch
Proposed
(12/1/2012)
Proposed
(12/1/2012)
Proposed
(12/1/2012)
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Hazard Analysis
Table 34 provides an overview of potential hazards to personnel through the
course of the project. Personnel hazards refer to potential harm incurred by any
individual. The development and implementation of the safety plan and protocols
ensure that these hazards are appropriately mitigated.
Risk
Laceration
Burns
Respiratory
Damage
Vision Damage
Allergic
Reaction
Sources
Knives,
routers,
saws, file,
Dremel tool
Chemicals
(FFFFg,
fiberglass
resin),
welders,
soldering
Iron
Chemicals
(epoxy,
solder),
fumes,
fiberglass
Welders,
fiberglass,
grinders,
projectile
debris
Epoxy,
chemicals,
fiberglass
Likelihood
Medium
Medium
Low
Consequence
Mitigation
Action
Serious injury
or death
Follow safety
protocols,
proper tool and
equipment use,
personal safety
attire, refer to
operators
manual
Discontinue
all
operations,
apply first
aid, contact
EMS
Minor to
serious injury
Brain damage
or death
Low
Partial to
complete
blindness
Low
Loss of
respiration,
inflammation
(Internal &
External)
Follow safety
protocols,
proper tool and
equipment use,
personal safety
attire, refer to
operators
manual
Follow safety
protocols,
proper tool and
equipment use,
personal safety
attire, consult
MSDS
Use of goggles,
force shields,
consult MSDS,
first aid kit
available, refer
to operators
manual
Use of gloves,
consult MSDS,
first aid kit
available
Discontinue
all
operations,
apply first
aid, contact
EMS
Discontinue
all
operations,
apply first
aid, contact
EMS
Discontinue
all
operations,
apply first
aid, contact
EMS, use
eyewash
Discontinue
all
operations,
apply first
aid, contact
EMS,
administer
antihistamin
es, safety
shower
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Hearing
Damage
Dismembermen
t
FFFFg,
Grinders,
Ignition,
Routers
Low
Projectiles,
Saws,
Launches
Low
Partial to
complete
deafness
Ear muffs,
consult MSDS,
first aid kit
available, refer
to operators
manual
Permanent
injury or death
Make sure
proper safety
measures are
taken,
operators
manual
Discontinue
all
operations,
apply first
aid, contact
EMS
Discontinue
all
operations,
apply first
aid, and
contact
EMS,
tourniquet
Table 34 Potential Hazards to Personnel
The material safety data sheet (MSDS) that the manufacturer provides contains
pertinent information about the material in consideration. It is comprised of sixteen
categories: identification, hazard(s) identification, composition/information on
ingredients, first-aid measures, fire-fighting measures, accidental release
measures, handling and storage, exposure controls/protection, physical and
chemical properties, stability and reactivity, toxicological information, ecological
information, disposal information, transport information, and regulatory information.
MSDSs are referred to when a hazard occurs in order to enact the most effective
mitigation. All team members shall be knowledgeable of the MSDS associated with
each hazardous material.
Operator manuals for each tool will be consistently referenced prior to each tool’s
usage. This ensures each tool is used as intended. According to the safety plan,
operator manuals for each tool used during the project will accompany the MSDSs
in the safety binder. These documents will be made available by the safety officer
at any location in which construction, testing, or launching of the vehicle could
occur.
It is important for all team members to be thoroughly briefed on the project risks,
FAA laws and regulations regarding the use of airspace, and the NAR high-power
safety code.
The team is aware that the FAA must be notified of planned launch activities. For
educational outreach events, notification to the closest airport within 5 miles of the
launch site is required 72 hours prior to launch. For subscale launches, flight
waivers are required to be obtained at least 45 days prior to the proposed activity.
For full scale launches, access to launch sites provided by West Texas Rocketry
Tripoli number 121 will be utilized.
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The National Association of Rocketry and Tripoli are recognized as the primary
rocketry associations of the United States. As such, their standards establish
precedence throughout high powered model rocketry. Along with these standards,
the team is cognizant of all federal, state, and local laws regarding unmanned
rocket launches and motor handling, including the following regulations:
CFR 101, Subchapter F, Subpart C: Amateur Rockets (Located in Appendix D)
CFR Part 55: Commerce in Explosives (Located Appendix D)
Handling and Use of Low-explosives Ammonium Perchlorate Rocket Motors
(APCP) (Located in Appendix I.12)
NAR Model Rocket Safety Code (Located in Appendix F)
Hazardous Waste Management (Located in Appendix E)
Fire Safety (Located in Appendix G)
Lab Safety (Located in Appendix H)
Table 35 provides a summary of legal risks that could occur during the course of
the project.
Risk
FAA Violations
NAR/TRA
Violations
Damage of
Property
OSHA
Violations
Personal
Injury
Likelihood
Severity
Consequence
Mitigation
Low
High
Legal
Repercussions
Adhering to
Regulations
Low
High
Legal
Repercussions
Adhering to
Regulations
Low
High
Legal
Repercussions
Insurance
Low
High
Legal
Repercussions
Adhering to
Regulations
High
Legal
Repercussions
Redundant
Calculations
and Safety
Preparedness
Low
Table 35 Summary of Legal Risks
Environment
Environmental effects of the project
In the event of an unrecoverable or damaged rocket, certain materials could be left
exposed to the environment. The biodegradability of each material used affects the
impact on the surrounding ecosystem. Much of the information on the hazards
posed to the environment and ecology is available in the individual MSDSs. The
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effects of materials used in the construction and launch of the rocket are
summarized in the following table 36.
Material
Ammonium
Perchlorate
Black Powder
Epoxy 2M DP420
Oil-Based Spray
Paint:
Clear Acrylic:
Prevalence
Mode of
Biodegradability
Motor propellant
Highly water soluble
Ejection charges
Fiberglass
connections, sealed
couplings
External fiberglass
structural
components
Remains solid
Iodization of local water
table
No known impact
Decomposition begins
within fifteen months
Leaching to local water
table
Soluble
Leaching to local water
table
Payload bay
Soluble
Structural
components
Remains solid
Deployment bag
Remains solid
Motor tube, rivets,
battery casing,
Highly water soluble
Recovery wadding;
Remains solid
Fiberglass
Nomex
Aluminum
Cellulose
Steel
Copper:
Sulfuric Acid
Kevlar
Silicon
Rip-Stop Nylon:
Impact on Environment
Leaching to local water
table
No known
environmental impact;
may pose ecological
hazard
No known
environmental impact;
may pose ecological
hazard
Long-term degradation
products
Long-term degradation
products
Attachment
hardware, ballast
system;
Avionics bay lining,
e-match lead wires
Highly reactive in air or
moisture
Batteries
Highly water soluble
Shock harnesses
Parachutes,
batteries
Parachutes, shear
pins
Remains solid
Long-term degradation
products
Long-term degradation
products; may pose
ecological hazard
No known impact
Reactive in air or moisture
Irritating vapors form
Remains solid
No known impact
Remains solid
Leaching to local water
table
Table 36 Effects of Materials used in Construction and Launch
Environmental effect on the project
While some aspects of the project may adversely affect the surrounding
environment, the environment can also have an impact upon the project. As
primary test launches will take place in Texas, during winter and spring, inclement
weather will likely fall on test dates. In response to unforeseen issues in the
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weather, alternate test dates have been scheduled and can be viewed in the
testing timeline in Figure 62. Environmental factors such as surrounding flora and
fauna or sedimentary projections could cause the launch vehicle to become
unrecoverable. These risks are outlined in Table 37.
Risk
Poor weather
Environment
prevents
recovery
Likelihood
Severity
Consequence
Mitigation
High
High
Delay in testing
Multiple test
dates
Medium
Medium
Possible loss of
Vehicle
Survey launch
site, recovery
tools
Low
High
Delay in testing
Multiple test
locations
Burn ban in
effect
Table 37 Environmental Factors
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IV) Payload Criteria
Selection, Design, and Verification of Payload Experiment
System Level Review
At a system level, the payload design fulfills the requirements of the SMD payload.
This payload option was selected because of the high difficulty level. The Tarleton
Aeronautical Team is honored to be able to apply personal skills and knowledge to
a real world scientific problem. The payload consists of redundant sensors and
power supplies. The payload electronics sub-team will implement the incremental
development process. At the completion of each increment, a fully functioning
product is developed. The deadline for the first increment is November 30th. The
deadline for the second increment is December 30th. During the first increment the
electronics are mounted on vertical aluminum rails. Connections between
electronics are assisted by perforated circuit boards. A model of the first increment
product is shown in Figure 30. After prototyping is successful, the next increment
of the development process includes fully integrating printed circuit boards (PCBs).
This reduces weight and size requirements of the payload.
Figure 30 Payload Design Configuration
The payload gathers measurements of pressure, temperature, relative humidity,
solar irradiance, and ultraviolent radiation (Requirement 3.1.3.1). Several readily
available sensors, along with a microcontroller equipped with the appropriate
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number of digital and analog inputs are selected. Redundant barometric pressure
sensors, redundant temperature sensors, redundant humidity sensors, redundant
UV radiation sensors, and redundant pyranometers are included in the design.
Measuring UV radiation and solar irradiance continuously throughout the flight is
achieved by placing a UV sensors and pyranometers facing every ninety degrees.
This design is shown in Figure 30. For consistency in data and sophistication in
design purposes the plan is to research an autonomous solar tracking system for
the UV radiation sensor and the pyranometer. The chosen microcontroller contains
the necessary data busses to communicate with all sensors.
The chosen microcontroller utilizes a microprocessor with a processing clock
speed capable of handling all the necessary sensors at the required frequencies
(Requirements 3.1.3.2 and 3.1.3.3). The clock speed of the microcontroller’s
microprocessor is 16 megahertz. Surface data collection terminates ten minutes
after landing (Requirement 3.1.3.4). The microcontroller is able to record the time
in milliseconds since the start of the most current software program. The flight
software will use sensor data to detect landing, note the current time in
milliseconds since the program started, then compare the current time to the time
at landing for every loop cycle; when the current time is 600,000 milliseconds (ten
minutes) larger than the time at landing, surface data collection will terminate.
The payload includes the ability to take two pictures during descent and three
pictures after landing, while keeping the orientation of the pictures with the sky
towards the top and the ground towards the bottom of the frame (Requirements
3.1.3.5 and 3.1.3.6). To ensure this functionality, our payload contains a camera
with a dedicated microcontroller, data storage device, and power supply. Several
options for camera orientation are considered. One option is to mount the camera
on a pendulum connected to a ball bearing. This method utilizes gravity and was
very efficient due to no power requirements. However, considering the orientation
of the payload, oscillations and rotational motion on descent is a major setback to
this option. The pendulum also has to be kept extremely stable and limited in its
range of motion; otherwise the camera could swing into the wall of the payload
housing.
Another alternative is to place a camera on each side of the payload, opposed 180
degrees. Throughout the flight a multitude of pictures would be taken, and during
post flight analysis the orientation of the pictures would be adjusted. This design
follows the Keep It Simple (KIS) principle; however, the team wants a more reliable
and sophisticated solution. The current design has the camera mounted on servo
motors. The correction angle of each servo is determined by an accelerometer
measuring the tilt of the payload. By measuring the change in orientation of the
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payload, the servo motors can correct the camera’s position such that pictures are
captured in proper orientation. Although this design requires a large power supply
and an additional sensor, the overall sophistication and reliability of the design are
the determining factors for selection of this option. Figure 31 demonstrates the
physical layout of the self-leveling camera.
Figure 31 Self-leveling Camera Configuration
Payload data is stored onboard and transmitted wirelessly to the ground station
(Requirement 3.1.3.7). A 16 gigabyte high speed data storage medium is selected.
Each measurement taken by the sensors along with any other usable data is
saved to this data storage medium. A 900 megahertz wireless transmitter
incorporates into the payload electronics. The microcontroller assembles a
telemetry string of sensor data. The sensor data is delimited by space characters.
The ground station includes a wireless receiver connected to a handheld high gain
Yagi antenna. A MATLAB graphical user interface program parses the telemetry
string, displays the readings, and saves each telemetry string to a file.
The payload carries a GPS tracking unit (Requirement 3.1.3.9). The GPS outputs
a National Marine Electronics Association (NMEA) Global Positioning System
Fixed Data (GGA) string at a rate of five Hertz. Each NMEA string is stored
onboard the payload. Also, the NMEA string is parsed for latitude, longitude, and
mean sea level altitude. The parsed values are transmitted to the ground station.
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The payload contains a liquid crystal display (LCD) screen. The LCD screen
displays a color image. The LCD screen assists in the pre-launch checklist phase.
The LCD screen displays sensor readings and the remaining power supply. The
LCD screen mounts inside the payload such that an outside observer is able to
clearly read the display. The LCD screen utilizes a dedicated power supply. The
LCD screen is powered down a short time prior to launch.
The payload includes a video camera. The video camera stores footage onboard
the payload and is useful for educational outreach and public relations after
launches. The video camera records both video and audio. The camera is an
independent electrical system from the other payload electronics; therefore the
video camera has a dedicated power supply. The video camera faces horizontally,
and due to the clear acrylic payload housing, is able to record video from within the
payload housing. Light emitting diodes ensures the camera is functioning properly
prior to launch.
The official scoring altimeter is located within the payload section of the vehicle.
The altimeter is mounted such that the scoring official can mark the device as the
official altimeter (Requirement 1.2.2.1). The altimeter mounts in a secure manner
to protect the sensor throughout the entire flight. The altimeter detects liftoff,
records apogee, and reports the official altitude post-flight through a series of
beeps (Requirement 1.2.1). The official altimeter is commercially available
(Requirement 1.2). The altimeter is on a separate circuit from other payload
electronics; therefore, the altimeter utilizes a dedicated power supply. In the
original proposed design, the official altimeter is located in the same compartment
as the parachute deployment altimeters. The decision to move the official altimeter
to the payload is made based on the fact that when the black powder charges fire
there is a chance that the spike in pressure could leak into the recovery altimeter
compartment. This pressure spike jeopardizes the accuracy of the apogee altitude
measurement. The clear payload housing also allows visual inspection of the
official altimeter.
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Required Subsystems
Atmospheric Data Gathering
Microcontroller
The atmospheric sensors send readings to an Arduino Mega 2560-R3
microcontroller, depicted in figure 32.
Figure 32 Arduino Mega 2560-R3 Microcontroller
The microcontroller includes an ATmega2560 microprocessor which has 16
analog input ports, 14 digital pulse width modulation enabled output ports, and 54
general purpose digital input and output ports. The ATmega2560 microprocessor
is capable of supporting at least four serial data interface devices, communicating
with Inter-Integrated Circuit (I2C) data interface devices, and containing a Serial
Peripheral Interface (SPI) bus for communicating with SPI data interface devices.
The microcontroller has built-in 3.3 volt and five volt voltage regulators. The
preceding characteristics are some of the determining factors to the
microcontroller selection. The manufacturer’s data sheet for the Arduino 2560-R3
can be found in Appendix J.1.
Alternative microcontrollers are considerations. The major alternative to the
Arduino Mega is the mbed NXP LPC1768. One desirable characteristic of the
mbed microcontroller is the faster clock speed. The mbed incorporates a 100
megahertz clock speed. However, the mbed compiler is not as user friendly as the
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Arduino. Also, the open source nature of the Arduino allows for more example
code and user advice. Another consideration is the Stellaris LM4F120 LaunchPad
created by Texas Instruments. This microcontroller is relatively new. Again, the
high clock speed is a desirable characteristic. The Stellaris Launchpad clock
speed is 80 megahertz. Because the Stellaris Launchpad is a new product, backorders and shipping time create a problem with using this product. The arrival date
for this component is in mid-December, which is long after development begins.
Table 38 lists the top four candidates for microcontroller selection. In addition to
this information, the availability, size, power consumption, and cost are
considerations throughout the selection process. The Arduino 2560-R3 is the
microcontroller selection.
Part Name
Distributor
Clock
Speed
Arduino 2560-R3
mbed NXP
LPC1768
Sparkfun
16MHz
Dimensions
(LxW)
2.125" x
4.3125"
Sparkfun
100MHz
1" x 2"
Arduino Uno- R3
Stellaris LM4F120
LaunchPad
Sparkfun
Texas
Instruments
16MHz
2.32" x 2.95"
80MHz
2" x 2.25"
Mass
2.3oz
12oz
1.06o
z
12oz
Input
Voltage
7 - 12V
4.5V 9V
Current
Draw
20 200mA
100 200mA
7 - 12V
4.75 5.25V
50mA
23 300mA
Cost
$58.95
$59.95
$29.95
Table 38 Top Four Candidates for Microcontroller Selection
Data Storage Device
The payload saves all sensor data to a 16 gigabyte high capacity micro secure
digital (SD) card. The micro SD card organizes data into a 32 bit file allocation
table (FAT32) file system. A micro SD card allows for quick retrieval of data.
Another consideration for data storage is Electrically Erasable Programmable
Read-Only Memory (EEPROM). However, retrieval of data from EEPROM is not
as convenient. Also, EEPROM has a lower data storage capacity. Another
consideration is the ENV-32X Embedded Data Logger. While the ENV-32X has a
very user-friendly interface, it has a limited storage capacity. Finally, a standard
size SD card is a possibility, but in this case the micro SD card’s smaller size is the
determining factor. Table 39 lists the top four choices for data storage medium.
Part Name
Part Number
Distributor
Capacity
micro SDHC
3FMUSD16FB-R
Walmart
16 GByte
$9.99
I2C EEPROM
525
Sparkfun
256 Kbit
$1.95
ENV-32X
11196
Sparkfun
32 Mbit
$34.95
SDHC
3FMSD32GBC10-R
Walmart
32 GByte
$17.99
Table 39 Top Four Choices for Data Storage Medium
Cost
$4.99
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Upon selection of the micro SD card for data storage, a means of interfacing the
micro SD card with the Arduino Mega 2560 is the next necessary determination in
the design process. Several distributors manufacture breakout boards specifically
designed to allow microcontrollers to easily communicate with micro SD cards over
the SPI interface. The best-fit micro SD breakout board adapter is distributed by
Adafruit. The board contains a built in five volt to three volt regulator. There are
light-emitting diodes (LED) to indicate when the micro SD card is read from or
written to. There is also a convenient locking mechanism to secure the micro SD
card and allow for easy removal of the card. The board utilizes the SPI data
interface to communicate with Arduino Mega 2560. Figure 33 shows the Adafruit
micro SD adapter.
Figure 33 Adafruit Micro SD Adapter
Table 40 is a trade and selection table listing the top three candidates for micro SD
adapter selection.
Part
Number
Distributor
Data
Interface
Dimensions
(LxW)
Mass
Input
Voltage
Current
Draw
254
Adafruit
SPI
1.25" x 1"
0.12oz
3 - 5V
150mA (max)
$15.00
32312
Parallax
SPI
1.11" x 1"
0.11oz
3.3V
0.5mA
$14.99
Sparkfun
SPI
0.94" x 0.94"
0.09oz
5V
0.5mA
$9.95
544
Cost
Table 40 Micro SD Adapter Selection
Pressure
In order to gather atmospheric pressure data, multiple barometric pressure
sensors are being researched. For redundancy, two separate barometric pressure
sensors are onboard the payload. For prototyping purposes, each sensor comes
preinstalled on a breakout board. The primary selection is a BOSCH BMP180. The
BMP180 uses a piezo-resistive sensor to detect applied pressure at a relative
accuracy of plus or minus 0.017 pounds per square inch (psi). The piezo-resistive
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sensor outputs an analog value, which converts to a 16 bit digital value by way of
an analog to digital converter. The BMP180 communicates with the Arduino Mega
2560 over the I2C data bus. The BMP180 data sheet can be found in Appendix
J.2. Figure 34 is a picture of the BMP180 installed on a breakout board.
Figure 34 BMP180 Breakout Board
The redundant barometric pressure sensor selection is the MS5611-01BA03 by
Measurement Specialties. The MS5611-01BA03 also uses a piezo-resistive
sensor. The relative accuracy is plus or minus 0.029 pounds per square inch.
While the BMP180 utilizes a 16 bit analog to digital converter, the MS561101BA03 implements a 24 bit analog to digital converter. It communicates with the
Arduino Mega using the I2C data interface or the SPI data interface. The MS561101BA03 datasheet is in Appendix J.3. Figure 35 is a picture of the MS561101BA03 installed on a breakout board.
Figure 35 MS5611-01BA03 Breakout Board
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Additional barometric pressure sensors are considered for selection. The
determining characteristics are availability, data interface compatibility, precision of
measurement, power consumption, and price. Table 41 is a pressure sensor trade
and selection table.
Part
Name
BMP180
MS561101BA03
Manufacturer
MPL115A
1
BOSCH
Measurement
Specialties
Freescale
Semiconductor
Inc.
BMP085
BOSCH
Data
Interface
Precision of
Measurement
I2C
±0.017psi
CSG Shop
SPI - I2C
±0.029psi
CSG Shop
SPI
±0.145psi
Sparkfun
I2C
±0.435psi
Distributor
DSS
Circuits
Input
Voltage
1.8 3.6V
1.8 3.6V
Current
Draw
32µA
(max)
12.5µA
(max)
2.375 5.5V
1.8 3.6V
10µA
(max)
12µA
(max)
Cost
$15.00
$29.99
$11.99
$19.95
Table 41 Pressure Sensor Trade and Selection
Temperature
Temperature data is gathered by the same sensors gathering pressure data. The
same piezo-resistive sensor is able to measure pressure and temperature.
Therefore, because there are redundant barometric pressure sensors, there are
also redundant temperature sensors onboard the payload. The purpose of this
decision is to maximize efficiency of the payload electronics. In contrast, an
individual sensor for the sole purpose of measuring temperature increases power
consumption and decreases room for other electronics in the payload. Table 42 is
a trade and selection table for the top considerations for temperature sensor
selection.
Part Name
Manufacturer
BMP180
MS561101BA03
BOSCH
Measurement
Specialties
TMP102
SHT11
Distributor
DSS
Circuits
Precision of
Measurement
±1.8° F
CSG Shop
±1.44° F
Texas Instruments
Sparkfun
±1.8° F
Parallax
Adafruit
±3.6° F
Table 42 Top Considerations for Temperature Sensor Selection
Cost
$15.0
0
$29.9
9
$5.95
$35.0
0
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Relative Humidity
The payload consists of two redundant humidity sensors. The first selection is the
Honeywell HIH4030. The HIH4030 is an analog sensor with no ADC. The analog
output transmits to an analog input pin of the Arduino Mega 2560. The HIH4030
sensor uses a thermoset polymer capacitive sensing element to measure relative
humidity to an accuracy of plus or minus 3.6 percent. For prototyping purposes,
Sparkfun retails a sensor preinstalled on a breakout board that meets the
necessary specifications. The sensor datasheet can be found in Appendix J.4.
Figure 36 is a picture of the HIH4030 installed on a breakout board.
Figure 36 HIH4030 Breakout Board
The redundant relative humidity sensor is the HH10D Hope RF relative humidity
sensor. In HH10D is also a capacitive type humidity sensor. The HH10D contains
two humidity measurement gathering chambers. The two measurements save to
EEPROM installed on the board. For precision purposes the two readings are
averaged to obtain accurate data. The HH10D communicates with the Arduino
Mega 2560 by means of a single digital frequency output. The sensor also utilizes
the I2C data interface to relay an initial calibration value. The HH10D datasheet is
in Appendix J.5. Figure 37 is a picture of the HH10D.
Figure 37 HH10D Breakout Board
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In addition to the two chosen humidity sensors, alternative sensors are
researched. Availability, data communication compatibility, data precision, power
consumption, and price are all determining factors throughout the selection
process. Table 43 is a trade and selection table of humidity sensors.
Part
Name
HIH403
0
Manufacturer
Distributor
Data
Interface
Precision of
Measurement
Input
Voltage
Honeywell
Sparkfun
Analog
±3.6% RH
HH10D
Hope RF
Sparkfun
I2C
±3% RH
SHT11
HIH613
0
Sensirion
Adafruit
Serial
±3% RH
Sparkfun
I2C
±4% RH
5V
2.7 3.3V
2.4 5.5V
2.3 5.5V
Honeywell
Current
Draw
200 500μA
Cost
$16.95
150μA
$9.95
90μW
1mA
(max)
$35.00
$29.95
Table 43 Humidity Sensors Trade and Selection
Solar Irradiance
The device that measures solar irradiance is referred to as a pyranometer. The
current payload design consists of redundant pyranometers. The payload consists
of two of each chosen pyranometer. Each pyranometer has a field of view of only
180 degrees; therefore they are opposed 180 degrees to each other. The first
pyranometer selection is the Apogee Instruments SP-110. The SP-110 measures
the spectral light range from 380 to 1120 nanometers. The sensor outputs an
analog value from zero to 350 millivolts. An increase of one millivolt corresponds to
a radiation increase of 0.456 Watts per square foot. The analog output connects to
an analog input pin of the Arduino Mega 2560. The SP-110 datasheet is located in
Appendix J.6. Figure 38 is a picture of the SP-110.
Figure 38 SP-110
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The chosen redundant pyranometer is a TAOS TSL2561. The TSL2561 measures
the spectral range from 300 to 1100 nanometers. The sensor includes two
photodiodes. The two photodiodes output an analog value which is then converts
to digital by onboard ADCs. Therefore, in contrast to the SP-110, the TSL2561
outputs a 16 bit digital value. As stated in the TSL2561 datasheet, a digital output
minimizes noise interference. The sensor communicates with the Arduino Mega
2560 over the I2C data interface bus. Through the use of the I2C interface, each
individual photodiode is read separately. While the sensor’s formal use is an
ambient light sensor, as stated in the datasheet, the raw output of the sensor is
irradiance which is useful to calculate ambient light. The TAOS TSL2561
datasheet is located in Appendix J.7. Figure 39 is a picture of the sensor
preinstalled on a breakout board.
Figure 39 TAOS TSL2561 Breakout Board
An alternative pyranometer is selected and will be tested to determine
performance characteristics. Determining factors in the pyranometer selection
process are availability, data communication compatibility, spectral range of
measurement, power consumption, and price. Table 44 is a trade and selection
table pyranometers.
Part
Name
SP-110
TSL2561
TEMT600
0
Distributor
Apogee
Instruments
Adafruit
Sparkfun
Data
Interface
Analog
I2C
Analog
Spectral
Range
380 1120nm
300 1100nm
360 970nm
Input
Voltage
Current
Draw
0V
0V
2.7 - 3.6V
0.5mA
6V
20mA
Table 44 Pyranometers Trade and Selection
Dimensions
(LxWxH)
0.94" x 0.94"
x 1.08"
0.75" x 0.75"
0.375" x
0.375"
Cost
$169.00
$12.50
$4.95
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Ultraviolent Radiation
Ultraviolet (UV) radiation is the light spectrum between ten nanometers and 400
nanometers. This range of the light spectrum is separated into three categories,
UVA, UVB, and UVC. The payload measures these three categories which include
the range of the light spectrum from 210 nanometers to 400 nanometers. When
selecting a sensor to gather UV radiation data, a deciding characteristic is the
range of the light spectrum each sensor measures. The payload includes
redundant UV sensors that measure the total UVA, UVB, and UVC spectrums.
Each sensor has a field of view of 180 degrees; therefore it is necessary to place
the sensors opposing 180 degrees to each other. The first UV sensor selection is
the Apogee Instruments SU-100. The SU-100 measures the light spectrum from
250 nanometers to 400 nanometers. A protective dome houses the sensor. The
SU-100 requires no voltage source, but rather accumulates power from the sun.
The SU-100 datasheet is in Appendix J.8. Figure 40 is a picture of the SU-100.
Figure 40 SU-100
The redundant UV sensor selection is the sglux TOCON_ABC3. The
TOCON_ABC3 is a preamplified UV sensor. The sensor measures the UV light
spectrum from 210 nanometers to 380 nanometers. The sensor outputs an analog
voltage representing UV light measurements. The analog output connects to an
analog input of the Arduino Mega 2560. In contrast to the SU-100, the
TOCON_ABC3 requires an input voltage of at least five volts. The TOCON_ABC3
datasheet is in Appendix J.9. Figure 41 is a picture of the TOCON_ABC3.
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Figure 41 TOCON_ABC3
Alternative UV sensors are products that are past considerations. The main
alternative to the two chosen sensors is the Solar Light Inc. PMA1107. A negative
feature of the PMA1107 is the high price. When choosing UV sensors, the
determining characteristics are availability, spectral measurement range, power
consumption, size, and price. Table 45 is a trade and selection table of the three
considerations for UV sensor.
Part Name
SU-100
TOCON_ABC
3
PMA1107
Distributor
Apogee
Instruments
sglux
Solar Light
Inc.
Spectral
Range
250 400nm
210 380nm
260 400nm
Input
Voltage
Current
Draw
0V
0A
Dimensions
(LxWxH)
0.925" x 0.925"
x 1.08"
2.5 - 15V
0.8mA
0.4" x 0.4" x 0.3"
$148.00
5 - 12V
1mA
1.6" x 1.6" x 1.8"
$525.00
Cost
$159.00
Table 45 UV Sensor Trade and Selection
Global Positioning System
The chosen GPS is the Locosys LS20031. The LS20031 contains a GPS antenna,
a MC-1513 GPS module, and a transistor-transistor logic (TTL) data interface. The
GPS has a refresh rate range of one Hertz to ten Hertz. The initial testing refresh
rate is five Hertz. Through testing, the optimal refresh rate is to be determined. The
GPS is also capable of receiving data from 66 different channels. Again, the
optimal channel is to be determined through testing. The GPS is accurate to within
nine feet. The GPS receives a NMEA string. There are six possibilities for NMEA
output message selection: Global positioning system fixed data (GGA), geographic
position – latitude/longitude (GLL), GNSS and active satellites (GSA), GNSS
satellites in view (GSV), recommended minimum specific GNSS data (RMC), and
course over ground and ground speed (VTG). The research concerning the GGA
output message which contains the following data: coordinate universal time,
latitude, longitude, number of satellites used, horizontal dilution of precision, mean
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sea level altitude, and geoid separation. The LS20031 datasheet is in Appendix
J.10. Figure 42 is a picture of the LS20031.
Figure 42 LS20031
Alternative GPS modules are available to conduct research. Although other GPS
modules have similar characteristics, the LS20031 is the best-fit for the project;
because, multiple members of the team have experience working with the module.
Also, availability, data communication compatibility, refresh rate, power
consumption, and price are determining factors for the decision. Table 46 is a
trade and selection table of the top three GPS modules that are considerations for
this type of a project.
Part Name
Distributor
Interface
Refresh
Rate
LS20031
Locosys
Serial
1- 10Hz
MTK3339
UBlOX
MAX-6Q
Adafruit
Serial
I2C,
Serial
1 - 10Hz
CSG Shop
5Hz
Mass
0.49o
z
0.3oz
0.33o
z
Dimensions
(LxWxH)
Input
Voltage
Current
Draw
1.18" x 1.18"
1.0" x 1.35"
x 0.25"
3 - 4.2V
29mA
$60.00
3 - 5.5V
$39.95
1.77" x 0.59"
3.3V
25mA
100mA
(max)
Cost
Table 46 GPS Modules Trade and Selection
Wireless Transmitter
The wireless transmitter that relays atmospheric sensor readings and GPS data is
the XBee-PRO XSC S3B. The XBee-PRO XSC S3B is capable of transmitting
telemetry to a line of sight range of 28 miles when coupled with a high gain
antenna. The transmitter operates at the 900 Megahertz frequency band. The 900
Megahertz band splits into 12 different channels to help prevent interference from
other devices transmitting at this frequency. The programming of the transmitter
incorporates a 64 bit personal area network identification number. The XBee-PRO
XSC S3B requires a low voltage of only 2.4 volts to 3.6 volts. The XBee-PRO XSC
$69.99
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S3B datasheet is available in Appendix J.11. Figure 43 is a picture of the XBeePRO XSC S3B.
Figure 43 XBee-PRO XSC S3B
The XBee brand is the only major consideration for wireless transmission in the
group’s view; however, XBee offers products that transmit over two different
frequency bands. The chosen transmitter utilizes the 900 Megahertz band while
the alternative XBee consideration utilizes the 2.4 Gigahertz band. The 2.4
Gigahertz XBee has a significantly lower telemetry range; although the 900
Megahertz XBee has a lower data transmission rate. The 2.4 GHz XBee data
transmission rate is 250 kilobits per second and the 900 Megahertz XBee data
transmission rate is only 10-20 kilobits per second. The telemetry range outweighs
the data transmission rate in terms of the team’s priorities. Table 47 is a trade and
selection table consisting of the two considered XBee modules.
Part Name
XBee-PRO
XSC S3B
XBee-PRO
ZB
Distributor
Telemetry
Range
Frequency
Band
Digi
28 miles
2 miles
Digi
Input
Voltage
900MHz
Data
Rate
10 20Kbps
2.4GHz
250Kbps
2.7 - 3.6V
Table 47 XBee Module Trade and Selection
Autonomous Camera Orientation System
2.4 - 3.6V
Current
Draw
215mA
(max)
205 mA
(max)
Cost
$42.00
$28.00
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The autonomous camera orientation system consists of a microcontroller,
photographic camera, accelerometer, data storage device, servo motors, and a
dedicated power supply. The camera continuously takes pictures throughout the
entire flight and saves them to the data storage device. Upon deployment of the
main parachute, the autonomous camera orientation system activates. The
accelerometer relays attitude measurements of pitch and yaw to the
microcontroller. The microcontroller then calculates the appropriate correction
angles for two servo motors. The camera mounts to the servos. Five minutes after
landing, the autonomous camera orientation system deactivates.
Microcontroller
The microcontroller that determines the servo motors’ angle is the Arduino Pro
Mini 328. The Arduino Pro Mini 328 contains an ATmega328 microprocessor. The
ATmega328 has significantly fewer digital and analog ports then the ATmega2560.
Fewer ports allow for smaller overall size. The Arduino Pro Mini 328 is only 0.7
inches by 1.3 inches, and has a mass of only 0.07 ounces. The ATmega328 is
capable of communicating with devices using the SPI data interface, I2C data
interface, and the serial data interface. The Arduino Pro Mini 328 datasheet can is
accessible in Appendix J.12. Figure 44 is a picture of the microcontroller.
Figure 44 Arduino Pro Mini 328 Mircocontroller
Other considerations for alternative microcontrollers did not meet the standards
the team looked for. The mbed NXP LPC11U24 is the secondary consideration.
The higher clock speed was a desirable characteristic, but the mbed compiler is a
negative factor throughout the mbed NXP LPC11U24’s consideration. The
compiler is only accessible when Internet access is available. This constraint
seems unacceptable. Microcontroller selection centers on features including
availability, clock speed, size, power requirements, and price. Table 48 is a trade
and selection table of the seriously considered microcontrollers.
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Part Name
Arduino Pro Mini
328
mbed NXP
LPC11U24
Stellaris LM4F120
LaunchPad
Distributor
Clock
Speed
Dimensions
(LxW)
Mass
Sparkfun
16MHz
0.7" x 1.3"
0.07oz
Sparkfun
Texas
Instruments
48MHz
2.13" x 1.02"
1 - 2oz
80MHz
2" x 2.25"
1 - 2oz
Input
Voltage
Current
Draw
5 - 12V
2.4 3.3V
4.75 5.25V
150mA
$18.95
50mA
23 300mA
$44.95
Cost
Table 48 Microcontrollers Trade and Selection
Camera
The selected photographic camera is the VC0706. Adafruit distributes a breakout
board with a preinstalled module. The camera sends each photograph to the
Arduino Pro Mini. The separate microcontroller is necessary due to the expected
delay time when saving the picture files to the micro SD. The VC0706
communicates with the Arduino using the serial data interface bus. The
photographs save to the micro SD card in Joint Photographic Experts Group
(JPEG) format. The camera mounts from the servo motors onto the breakout
board using the prefabricated mounting holes. The VC0706 datasheet can is
accessible in Appendix J.13. Figure 45 is a picture of the camera.
Figure 45 VC0706 Photographic Camera
Alternative camera selections are past considerations. The Sparkfun SEN-10061
is the first to attain status as a serious alternative. The major negative
characteristic of the SEN-10061 is the lower resolution of the pictures. There are
multiple figures available for the resolution of the SEN-10061. The distributor
website, Sparkfun.com, states that resolution is 640 pixels by 480 pixels. Table 49
is a trade and selection table contacting the top three seriously considered camera
modules.
$4.99
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Part
Name
VC0706
SEN10061
OV7670
Distributor
Data
Interface
Pixels
Adafruit
Serial
640 x 480
Sparkfun
Electro
Dragon
Serial
160 x 120
I2C
640 x 480
Dimensions
(LxWxH)
1.26" x 1.26" x
0.25"
1.26" x 1.26" x
0.25"
0.24" x 0.24" x
0.16"
Input
Voltage
Current
Draw
5V
75mA
$42.00
5V
0 - 100mA
$49.95
3.3V
60mW
Cost
Table 49 Camera Modules Trade and Selection
Servos
Servos control the orientation of the camera. One servo mounts to the aluminum
rail on the side of payload. The second servo mounts to the first servo’s rotational
arm. The camera mounts to the second servo’s rotational arm. This setup allows
for camera orientation in the horizontal and vertical directions. There are many
types of servo motors. The optimal servo for the payload has characteristics of
being small but not so small that the torque compromises the function as a result.
Both servo motors are HS-85BB+ Mighty Micros. The Mighty Micro is small,
lightweight, and strong. A pulse width modulated output signal from the Arduino
Pro Mini 328 controls the servo. The gears within the Mighty Micro are nylon.
Another possible material for gears is metal. Servo motors that contain metal
gears burn out if they become stuck. In contrast, when servo motors contain nylon
gears, usually the gears strip before the motor burns out. The Mighty Micro also
contains a ball bearing to make the motor move smoother. The HS-85BB+ Mighty
Micro datasheet is accessible in Appendix J.14. Figure 46 is a dimensional layout
of the servo.
Figure 46 Servo Dimensional Layout
$9.30
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Alternative servo motors are prior considerations of the team. The Standard Servo
f by Parallax is definitively too large. The HS-55 Sub-Micro is too small and weak.
Table 50 is a trade and selection table of the top three seriously considered servo
motors. The determining characteristics are availability, torque, size, power
requirements, and price.
Part Name
HS-85BB+
Mighty Micro
Standard Servo
(#900-0005)
HS-55 SubMicro
Distributor
ServoCity
Parallax
ServoCity
Torque
41.66
oz-in
Mass
38oz-in
15.27
oz-in
1.55oz
0.67oz
0.28oz
Dimensions
1.14" x 0.51" x
1.18"
2.2" x 0.8" x
1.6"
0.89" x 0.45" x
0.94"
Input
Voltage
Current
Draw
4.8V
240mA
$19.99
6V
190mA
$12.99
4.8V
150mA
$9.99
Cost
Table 50 Servo Motors Trade and Selection
Accelerometer
An accelerometer measures acceleration in multiple planes. Because gravity is
essentially constant, it is possible to accurately measure the tilt by relative
acceleration. The chosen accelerometer is an ADXL345. The ADXL345 measures
acceleration in the X, Y, and Z planes. The ADXL345 utilizes the I 2C data interface
bus. The X and Y measurements of the accelerometer represent the tilt and yaw of
the payload. Each reading stores to the micro SD card for post flight analysis. The
measurements also determine the angle of the servo motors. The ADXL345
datasheet is accessible in Appendix J.15. Figure 47 is a picture of the sensor
following installment on a breakout board.
Figure 47 ADXL345 Breakout Board
Alternative accelerometers are in consideration for the project. The most serious
alternative accelerometer is the Memsic Mx2125. The sensor is still in
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consideration. The Mx2125 utilizes a pulse width modulated output for each of the
axis measurements. Both the Mx2125 and the ADXL345 are facing head to head
testing by the Tarleton Aeronautical Team. Through testing, the most reliable
sensor remains to win its spot. The Mx2125 is only capable of measuring
acceleration of plus or minus 3Gs, and the ADXL345 is capable of measuring plus
or minus 16 Gs. This is the determining factor for the current selection. However,
the ADXL345 measures three planes of acceleration while the Mx2125 measures
only two. Two planes of measurement are all that is necessary for the camera
orientation system. Table 51 is a trade and selection table of the serious
competition among accelerometers.
Part
Name
Distributor
Interface
Range of
Measurement
Input
Voltage
Current
Draw
ADXL345
Sparkfun
I2C
3
±16 g
2 - 3.6V
40μA
$27.95
Mx2125
Parallax
PWM
2
±3 g
3.3 - 5V
5 mA
$29.99
ADXL335
Sparkfun
I2C
3
±3 g
1.8 - 3.6V
320μA
$24.95
Axes
Cost
Table 51 Accelerometers Trade and Selection
Video Capture
Video Camera
The payload houses a video camera. The chosen video camera is the VCC-003MUVI-BLK. The camera is extremely small, lightweight, and capable of recording
video for up to 90 minutes. The camera has a rechargeable battery preinstalled.
The camera utilizes a dedicated micro SD card. The selected micro SD card that
stores the video is the same type of micro SD selected to save sensor data. The
camera powers on manually very near to launch time. The camera is completely
independent from the other payload electronics. The camera mounts to the rails of
the payload. The resolution of the video is 640 pixels by 480 pixels. After the flight,
the micro SD card removes manually from the camera and the footage facilitates
educational engagement purposes and public relations. Table 52 is a datasheet of
the VCC-003-MUVI-BLK. Figure 48 is a picture of the selected video camera.
Figure 48 VCC-003-MUVI-BLK Video Camera
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Research on alternative video cameras led to this selection. The most competitive
video camera is the FlyCamOne2. The FlyCamOne2 has the same capacity for
length of video. However, all distributors of FlyCamOne2 state that the product not
currently available. Table 50 is a trade and selection table of the top two
competing video cameras.
Part Name
VCC-003MUVI-BLK
Distributor
Frame
Rate
Mass
Adorama
20 - 25fps
1.76oz
FlyCamOne2
Hacktronics
25fps
1.31oz
Dimensions
(LxWxH)
1.10" x 2.17" x
0.79"
1.57" x 3.15" x
0.55"
Battery Usage
Time
Cost
70 - 90 minutes
$49.95
30 minutes
$69.95
Table 52 Top Two Video Cameras Trade and Selection
Liquid Crystal Display
Sparkfun distributes the selected LCD, part number 11062, which mounts on a
breakout board. The LCD has a color display with a resolution of 132 pixels by 132
pixels. The Arduino Mega 2560 microcontroller controls the LCD. The LCD screen
utilizes the SPI data interface bus. One desirable feature of the chosen LCD
screen is the low voltage requirements. At a minimum, the LCD screen requires
3.3 volts. The LCD-11062 datasheet is available in Appendix J.17. Figure 49 is a
picture of the chosen LCD screen.
Figure 49 Sparkfun LCD-11062 Screen
Alternative LCD screens are serious past considerations. A smaller black and
white LCD and a larger touch screen color LCD are the two most seriously
competitive alternatives. The black and white LCD falls short of some needed
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conventions; because, it is a desirable feature to have color rather than black and
white. The touch screen falls short; because, the functionality is not necessary.
Table 53 is a trade and selection table of the three most serious competitors for
LCD screens.
Part
Number
LCD11062
LCD10168
335
Distributor
Sparkfun
Sparkfun
Adafruit
Resolution
132 x 132
pixels
48 x 84
pixels
240 x 320
pixels
Dimension
s
Input
Voltage
Current Draw
3.3 - 6V
108 – 324mA
$34.95
No
1.5" x 2.5"
1.75" x
1.75"
3.3 - 6V
240 – 320mA
$9.95
Yes
2.5" x 3.2"
3.3 - 5V
150mA
Color
Yes
Cost
Table 53 LCD Screen Trade and Selection
Official Scoring Altimeter
The official scoring altimeter beeps the apogee height upon landing. The height
records in feet in accordance with the altimeter’s data. The chosen official scoring
altimeter is the Adept A1E. The altimeter wiring is completely independent of the
other payload electronics; therefore, the altimeter utilizes a dedicated power
supply. The required power supply is a twelve volt battery which comes with the
altimeter upon the receipt of the order. The battery that ships with the altimeter is
the GP-23A Alkaline Lighter Battery. This power supply allows the altimeter to
function for up to 10 hours. The altimeter datasheet is available in Appendix J.18.
Figure 50 is a picture of the Adept A1E.
Figure 50 Adept A1E
The Adept A1E is the strongest altimeter due to its small size and long battery life.
Alternative scoring altimeters offer strong competition. The alternatives are able to
fire black powder and record the apogee altitude. The PerfectFlite Stratologger is
one of the altimeters recovery deployment employs. The decision to select an
altimeter that is specifically designed for only reading out the apogee altitude in
$40.00
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beeps and not a multifunctional deployment altimeter fits the parameters best.
Similar reasoning was leads to the rejection of the alternative official scoring
altimeter. Table 54 is a trade and selection table of the three serious
considerations for official scoring altimeters.
Name
Distributor
Dimensions
Input Voltage
Cost
A1E
Adept
0.55" x 2.2"
12V
$29.95
Stratologger
PerfectFlite
2.75" x 0.9"
9V
$79.95
ARTS
Loki Research
3.75" x 1.4”
9 - 15V
$190.00
Table 54 Official Scoring Altimeters Trade and Selection
Power Supply
Eight nine volt batteries power the payload. This power supply provides more than
adequate voltage to power all the electronic components. Voltage regulators level
the voltage down to five volts and three volts, depending on requirements for each
component. The payload is able to function at full operation for at least two hours.
The power supply distributes power to the components so that if single batteries
fail the entire system will not fail.
Battery
Research and analysis lead to the battery selection in place. The determining
characteristic is capacity. The U9VLBP has a large capacity of 1.2 amp-hours.
This capacity exceeds the current requirements of the payload. Alternative
batteries are prior considerations, but both alternatives have a lower capacity.
However, both alternatives are more readily available. Table 55 is a trade and
selection table of the three most serious considerations for batteries.
Part Name
Manufacturer
Capacity
Voltage
U9VLBP
Ultralife
1.2 Ah
9V
$5.50
Coppertop Alkaline
Duracell
580 mAh
9V
$11.00
Advanced Lithium
Energizer
750 mAh
9V
$9.49
Table 55 Batteries Trade and Selection
Price
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Performance Characteristics
The performance of the payload system characterizes how well the system
completes the necessary objectives. The main objectives of the payload are to
store and transmit atmospheric and GPS data, to collect pictures at the proper
orientation, and to record flight footage. Research and analysis of many sensors
and electrical components view evaluation criteria for the success of the payload
system. Wiring diagrams and power budgets formulate and verify through analysis
and testing. Researching cameras and formulating and prototyping mechanisms
for leveling the camera throughout flight is a job necessary to sound design. Exact
dimensional CAD drawings and 3D renderings illustrate and help evaluate the
fitment of the camera orientation system. Verification of performance through
prototyping and flight testing is ongoing. Evaluations of video cameras and their
ability to record video footage are paramount for the objectives of the flight.
Testing the video camera lends itself to verification of the camera. Table 56
represents the evaluation and verification metrics for the payload subsystems.
Performance
Characteristics
Evaluation
Verification
Atmospheric Data
Gathering
Store and transmit
atmospheric and GPS data
Research and
analysis of
sensors and
electronics
Analysis of schematics
and testing of
components
Camera Orientation
Save pictures of the proper
orientation
Research and
CAD drawings
Prototyping and flight
testing
Video Footage
Store footage of the entire
flight
Research of video
cameras
Testing
Subsystem
Table 56 Payload Subsystems Evaluation and Verification Metrics
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Verification Plan
The verification plan reflects how each requirement to the payload system is in a
satisfactory state of completion. The SOW enumerates requirements that appear
in brevity in Table 57. The table also lists satisfying features of the design for the
requirement in question. Ultimately, design features require verification to ensure
that it actually meets the requirement. This goal is achievable through testing,
analysis, and/or inspection.
SOW
#
3.1
Requirement
Launch vehicle shall carry
a science or engineering
payload
Satisfying Design
Feature
Verification Method
SMD payload selection
Inspection, Analysis
3.1.3.1
Measurements of
pressure, temperature,
relative humidity, solar
irradiance and ultraviolet
radiation
An appropriate sensor
for each of these
measurements has been
selected
Testing, Analysis,
Inspection
3.1.3.2
Measurements occur
every 5 seconds during
descent
Main flight
computer/microcontroller
sampling rate
Testing, Analysis
3.1.3.3
Measurements occur
every minute after landing
Main flight
computer/microcontroller
sampling rate
Testing, Analysis
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Main flight
computer/microcontroller
shall halt all data
collection after 10
minutes
Testing, Analysis,
Inspection
3.1.3.4
Data collection will cease
10 minutes after landing
3.1.3.5
Payload will take 2
pictures during descent
and 3 after landing
Control of camera to
appropriate image
capture
Testing
3.1.3.6
Pictures taken must be
properly oriented such
that the sky is at the top
of the frame, ground at
the bottom
Self-leveling camera
mount
Testing, Inspection
3.1.3.7
Data from payload will be
stored onboard and
transmitted wirelessly
microSD onboard
storage; radio link
between Xbee radios
from rocket vehicle to
ground station
Testing
3.1.3.8
Separations of payload at
apogee is allowed, but
such separation may
cause drifting outside
recovery area
The payload does not
detach from the rocket
vehicle
Inspection
Locosys LS20031 is
incorporated in the
payload design
Inspection, Testing
(satisfying design
feature) experimental
logic, approach, and
method of investigation
Analysis
3.1.3.9
3.2
Payload will carry GPS
Data from payload will be
collected, analyzed, and
reported by team
following scientific
method
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3.5
Payload must be
designed to be
recoverable and reusable
Payload will be housed
in the main vehicle and
remain there during the
entirety of the flight
Testing, Inspection
Table 57 - SOW Verification
Preliminary Integration Plan
Payload Integration Plan
The payload integrates into the payload housing structure of the vehicle in a
simple fashion. A payload framework consisting of two bulkheads connected to
each other by two aluminum rails houses the payload. Since the payload housing
structure is part of the vehicle, the payload framework uses dimensions to ensure
compatibility between the payload and the vehicle.
Preparing the payload for flight is simple prior to payload framework integration
with the vehicle. Once the payload is ready for flight, the payload framework is
ready for integration. At this time, the framework inserts from the lower portion of
the payload housing structure. Once the payload framework is in place, it secures
with the use of two screws from the exterior of the payload housing structure into
the bulkhead. At this point, the physical integration of payload is complete.
Precision of instrumentation, repeatability of measurement, and
recovery system.
The instrumentation that takes measurements is commercially available on
breakout boards that manufacturers verify and test. Data sheets that
manufacturers provide for each sensor outline the precision and error tolerance.
Collected data compare to known data from sources such as NOAA to verify
accuracy and precision. The repeatability is easily accomplishable, since the
payload is modular and sensor data records to microSD cards. The only thing
necessary to gather a new set of data is the removal of the payload to recover the
data, install a fresh microSD card, and reload the payload into a fresh preparation
of the launch vehicle for a new flight. The parachutes consist of durable, siliconcoated ripstop nylon. This allows the parachutes to function repeatedly, requiring
only new black powder charges before preparing the recovery system for a new
flight. Table 58 lists the precision of each sensor in the payload.
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Purpose
Barometric Pressure
Barometric Pressure
Temperature
Temperature
Humidity
Humidity
Solar Irradiance
Solar Irradiance
Ultraviolent Radiation
Ultraviolent Radiation
GPS
Accelerometer
Official Altimeter
Product
BMP180
MS5611-01BA03
BMP180
MS5611-01BA03
HIH4030
HH10D
SP-110
TSL2561
SU-100
TOCON_ABC3
LS20031
ADXL345
Adept A1E
Table 58 Payload Sensors Precision
Precision
±0.017psi
±0.029psi
±1.8° F
±1.44° F
±3.6% RH
±3% RH
±5%
±5%
±10%
±10%
±9.84ft
±4.3mg
±1ft
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Drawings and Electrical Schematics
Figure 51 is a conceptual wiring diagram of the atmospheric data gathering
subsystem. Determination of proper wiring was comes from a careful review of the
datasheets.
Figure 51 Atmospheric Data Gathering Conceptual Wiring
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Figure 52 Camera Orientation Wiring
Electrical Schematic
Figure 53 represents the schematic for the payload sensors. The schematic is fully
developed to allow for the design and development of printed circuit boards (PCB).
Though the current payload design uses an array of breakout boards, a PCB
design would lead to a more compact and efficient payload system. Due to
resolution limitations and the large size of the perf board circuit, the figure cannot
properly display the labels of each component. To remedy this, a PDF is posted on
the Tarleton Aeronautical Team website to show the schematic in full resolution.
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Figure 53 PCB Schematic for Payload Sensors
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Figure 54 I^2C Sensors
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Figure 55 Radio Communications
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Figure 56 MicroSD Card Reader
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Figure 57 Analog Sensors
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Figure 58 USB Interface
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Figure 59 GPS Module
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Figure 60 Voltage Regulators
PCB Layout
During the course of testing and design, the payload circuitry progresses towards
a complete PCB layout. Figure 61 demonstrates the preliminary design for the
PCB layout. A number of advantages arise from using and creating a custom PCB
design. The size of the payload housing reduces drastically. Altogether the
breakout boards add to almost 20 in2. The design uses a 4.75” x 4.75” surface and
even then leaves a large amount of unused space and room for optimization. After
thorough planning, the PCB design can reasonably be reduced to nearly half the
size of a breakout board system.
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Figure 61 Preliminary PCB Layout
Furthermore, utilizing a custom-designed PCB layout optimizes the efficiency to
the designer’s exact specifications. Figure 62 displays a component listing printout
from the PCB design. Each component is specifically chosen to minimize the
amount of ESR introduced to the system. Copper wires and through-hole
components inherently have larger amounts of ESR. This causes parasitic
capacitance and resistances to effect the circuitry adversely. Each component in
figure is chosen for its SMT design. This smaller area of conductive material
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reduces the ESR drastically (from approximately 7Ω to 0.05Ω). The difference
between these two component constructions is available in related figures.
Figure 62 Component Listing
Figure 63 SMT (left) vs. Through Hole Devices (right)
Ultimately, the PCB design will be optimized over the course of the testing phase
to yield a fully-functional and extremely efficient payload circuit. The difficulty of
this PCB design and build process sets this payload design apart from others. To
create this preliminary design cost approximately 100 man-hours. To improve
upon this design is a tedious and challenging process; however, it results in an
exceptional payload circuit.
Cross-Component Compatibility
The atmospheric data gathering subsystem has a unidirectional data flow. The
flow of data travels from the atmospheric sensors relaying measurements to the
Arduino Mega 2560. Once sensor measurements come in, they transfer and save
to a file on the 16 gigabyte micro-SD card. Pre-flight measurements display on the
LCD screen just prior to launch throughout the flight. The Arduino Mega 2560
sends telemetry data to the XBee Pro XSC S2B at regular intervals. The XBee Pro
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XSC S2B wirelessly transmits the telemetry data to the ground station. Figure 54
portrays the flow of data for the atmospheric data gathering subsystem.
Figure 64 Atmospheric Data Gathering Subsystem Data Flow
The autonomous camera orientation subsystem also has a unidirectional flow of
data. The data flow begins with the ADXL345 accelerometer measuring the tilt and
yaw of the payload. The tilt and yaw data transfers to the Arduino Pro Mini. At
regular intervals, the VC0706 photographic camera takes pictures and transmits
them to the Arduino Pro Mini. The Arduino Pro Mini sends angular settings to the
HS85BB Mighty Micro servo motors in accordance with the accelerometer
readings while simultaneously saving pictures to the micro-SD card. Figure 55
models the data flow for the autonomous camera orientation system.
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Figure 65 Autonomous Camera Orientation System Data Flow
In order for the Arduino microcontrollers to properly function, the user must design
and write software that governs the desirable function. The software dictates a
protocol determining how each separate component communicates with the
microcontroller. Software design and analysis follows the object oriented
programming paradigm, allowing the code to easily update and debug. The
software development process begins with conceptual designs. Implementation
and deployment of code is the final phase of the process. This method generates
efficient and redundant software. The software development process is under way.
The first steps include the generation of flow charts, which function as high level
layouts for the software. There are two programmable microcontrollers onboard
the payload; therefore, two separate flow charts exist. Figure 66 is the flow chart
for the atmospheric data gathering subsystem and Figure 67 is the flow chart for
the autonomous camera orientation subsystem.
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Figure 66 Atmospheric Data Gathering Software Flow Chart
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Figure 67 Autonomous Camera Orientation Software Flow Chart
Payload Concept Features and Definition
Creativity and Originality
A clear acrylic section of the vehicle body houses the payload. This allows visual
inspection of all sensors and indicator LEDs along with verification readouts from
the LCD during assembly and pre-launch operations. Given the payload
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integration design, the payload is easily and quickly removable. Given the modular
design of the power circuits, disposable batteries are replaceable when necessary.
Although testing and prototyping begins with breakout boards, ultimately the goal
is to work towards designing the PCBs from scratch to size and wiring
specifications. These completely original PCBs and the designs are realized as a
final product. A silkscreen image prints a team (or university) emblem on the board
in Tarleton purple.
Figure 68 - PCB Board
The self-leveling camera system ensures that photographs are taken in proper
orientation as required. The tilt sensor detects any changes in orientation, where
two servo motors correct movement. The LCD displays appropriate data and
relevant readouts to ensure and verify payload functionality. The team plans to use
video camera for flight documentation and plans to use this as an educational
outreach tool.
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Figure 69 - Self-Leveling Camera System
Uniqueness and Significance
The scientific payload is significant in that it meets the SMD requirements set forth
in the NASA SLP SOW. All design features and component selections reflect
meeting the customer prescribed specifications. The ability of the payload to
gather atmospheric data is significant for the analysis of changing conditions as
the vehicle varies in altitude. All data compiling is useful for finding correlations to
altitude such that atmospheric conditions from ground level to vehicle apogee can
model and test for significance.
The clear acrylic airframe houses the payload. This is unique in that it allows visual
inspection of all components of the payload externally. Once preparation and
integration of the payload takes place in the vehicle, a final verification ensures
flight readiness relating to the payload. An LCD screen displays relevant checks
and indicators from the flight software.
The means to maintain proper orientation of the camera is unique as well.
Because the orientation of the payload is no guarantee upon descent, a corrective
measure uniquely takes place. The chosen solution creates and implements a selfleveling camera. Any changes to the orientation of the payload relative to starting
position at the launch pad can result in measurement and compensation for the
discrepancy.
A video camera in the payload captures video of vehicle flights. Although this is a
standard practice in rocketry, video recordings provide flight documentation that
can be useful in educational engagement and community outreach.
Suitable Level of Challenge
The design complexity and implementation of a functional SMD payload is an
extremely taxing challenge. Not only must appropriate components reside in the
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vehicle, they must orchestrate in a manner that achieves a useful atmospheric
measuring instrument. This involves proper interfacing, sufficient power supply,
and adequate programming logic of all constituent pieces of the payload.
Although very challenging, design of the electrical circuits ultimately seek PCB
implementation. This is a significant advantage over breakout board and
perforated board mounting. The space saving and overall efficiency of the PCB
design merits it as the fruition of majors aims of the payload electronics. Despite
the level of challenge and time necessary to complete this design, it is worthwhile
in view of the integrity of the payload system as a whole.
Science Value
Payload Objectives
The main objectives of the payload are to store and transmit atmospheric and GPS
data, capture photos of the correct orientation, and to record video footage of the
entire flight. The payload must receive sufficient power to operate while on the
launch pad, during flight, and after landing. The atmospheric data collection
illuminates the change in atmospheric variables depending on altitude. The GPS
data shows the flight path of the vehicle and aids in recovery. The camera
orientation system tests a multi-servo, autonomous orientation device. The video
footage is profitable for public outreach in the aftermath of the flight.
Payload Success Criteria
If the payload objectives succeed, the mission is a success. The atmospheric data
is in this case accurate and stores at a minimum of one reading from each sensor
every five seconds. Transmission of data from the payload to the ground station
operates throughout the entire flight. Two photos during descent and three after
landing capture valuable references in the flight. The photos should orient in such
that they portray the sky at the top of the frame and the ground towards the
bottom. The video footage spans the entire flight. Table 59 summarizes the
payload objectives, their science value and success criteria.
Payload Objective
Store and Transmit
Atmospheric and GPS Data
Science Value
Success Criteria
Represents Change in
Atmospheric Variables
Depending on Altitude
Accurate Data, Collected at Required
Frequency, Transmitted Throughout Flight
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Camera Orientation
Record Video Footage
Test Multi-Servo
Orientation Device
Two Pictures During Descent and Three
after Landing, Correct Orientation
Aid in Public Outreach
Record Entire Flight, Collect Usable
Footage
Table 59 Payload Objectives Summary
Experimental Logic, Approach, and Method of Investigation
The SMD payload gathers data from approximately 5280 feet above ground level
(AGL) down to the landing site, collecting altitude-varied data for five atmospheric
variables: pressure, temperature, relative humidity, solar irradiance, and ultraviolet
radiation. This data determines the accuracy of the payload sensors and the
statistical correlations between each of the various variables. Together, these two
calculations develop a regression model for each variable. By creating a model to
represent these correlation effects, a new and comprehensive formula could
demonstrate causal relationships between these five variables or any derivative
subset.
Regression Model
With a large number of samples ranging across the various test and demonstration
flights, the data plots determine the order of polynomial of best fit. If the order of
this polynomial is one then a simple linear regression model of the form can result
through statistical analysis of the data. If the order of the proposed polynomial
does not equal one, then local regression determines the most likely model to
represent the data and predict new values. This process is repeatable for each
variable against altitude and then in any combination of subject variables. Once
this model establishes itself and verifies, it can stand to test and improve any
functions relating altitude to any of the variables this study includes.
Correlations between SMD Sensor Readings
After establishing an acceptable model, the distributions of each variable aid in
finding covariance between any two variables and, consequently, the correlation
between those variables. After determining the correlation, the model can evolve
as necessary to provide more accuracy.
Accuracy of Sensors
Though the datasheets for each sensor propose a certain level of accuracy, this
cannot guarantee the sensors will perform to this level within the payload circuitry.
By comparing collected data against atmospheric measurements from national
databases, any discrepancy can establish itself. Furthermore, using both data sets
can establish confidence intervals to ensure new data are within a particular range
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of the presumably absolute readings from these databases. This gives a measure
of accuracy as it pertains specifically to the payload circuitry.
Test and Measurement, Variables, and Controls
The variables include altitude, temperature, atmospheric pressure, relative
humidity, solar irradiance, UV radiation and random electrical noise in the payload
circuitry. The
-terms in the example model above comprise of vectors of each individual payload
measurement. The –term’s resultant vector represents the unobservable electrical
noise in the system.
The controls in consideration are the measurements from national agencies, such
as the National Oceanic and Atmospheric Administration (NOAA), whose
measurement devices have presumably negligible errors.
The measurement process centers on data acquisition from the payload sensors.
The testing process occurs post-flight in the form of a statistical analysis of the
acquired data.
Relevance of Expected Data and Accuracy/Error Analysis
The findings compare against documented formulae for atmospheric
measurements to determine their validity. Furthermore, the models could
potentially represent undocumented relations between these variables.
The accuracy and error analysis are an inherent part of the entire process. All
errors factor into the final regression model. Small differences from the expectant
values will demonstrate as random noise in the model.
Preliminary Experiment Process Procedures
Numerous hypotheses exist prior to measurement and testing, but the testing of
these hypotheses still requires the analysis of future data. Thus, no actions in
experimentation are complete.
Safety and Environment (Payload)
REFER TO SAFTEY AND ENVIROMENT (VEHICLE)
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V) Project Plan
Budget Plan
A summary of the costs for various aspects of the project assembles into a chart
format for efficiency and ease of use. Tables of pertinent data follow.
Vehicle Budget
The table below lists cost by subsystem for one complete build out of the proposed
design. The chart of subsystem budgets is comprehensive such that all
subsystems necessary to mission completion are present with their respective
budget totals. Itemized budgets of each subsystem appear in Appendix A.
Subsystem
Vehicle
Recovery
Payload
Propulsion
Total
Cost
$765.22
$1383.35
$1558.24
$434.42
$4202.33
Table 60 Proposed Rocket Vehicle Budget Summary
Testing Budget
Testing is one of the most important areas in the development of any system,
particularly a vehicle of this nature. This fact is not lost to the future engineers at
Tarleton. The importance of testing a system is two-fold; it can eliminate product
flaws and potential failures, and it can prove and improve upon a system’s overall
safety. Taking the gravity of the task into consideration, the following budget was
summarizes a preliminary testing budget. The following chart includes preliminary
test budgets for each subsystem. Itemized test budgets are available in Appendix
A.
Subsystem
Vehicle
Recovery
Payload
Propulsion
Total
Unit Cost
# of Test
$765.22
$1383.35
$1558.24
$434.42
Table 61 Preliminary Testing Budget Summary
3.5
2
4
5
Unit Total
$2,678.27
$2,766.7
$6,232.96
$2,172.1
$13,850.03
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Outreach Budget
Outreach and community involvement events are critical to any engineering project
both for publicity as well as educational opportunities. The project hosts many
events that encourage students to participate in STEM fields. In order to host an
effective outreach program a budget has been generated. The following chart
includes a summarized budget of the outreach program.
Event
Cost
Community Events
Scholastic Events
Equipment
Travel
Total
$1,709.43
$719.60
$140.74
$1,100
$3,669.77
Table 62 Outreach Budget
Travel Budget
In order to meet all requirements of the SOW, the team must travel to Alabama
and present their project. The logistics of moving a large number of personnel
requires thought and careful planning in order to stay within budget. As of this
report, the number of personnel traveling to Alabama is 15 and the required
budget totals to $8200. The following chart consists of rough estimates of travel
costs.
Service
Hotel
Gas
Van Rental
Meals
Total
# of People
Est. Cost
15
15
15
15
$4,200
$1,000
$1,000
$2,000
$8,200
Table 63 Estimated Travel Budget
Budget Summary
The following chart includes the projected budget for completing the project. The
task of completing the NASA USLI is a complex interdisciplinary endeavor that
tests the team’s knowledge and skills, including management of a budget. The first
step in managing a budget is devising such a budget that is sufficient in meeting all
costs necessary to complete the mission. Table 64 breaks down the known project
costs, and a detailed budget is available for review in Appendix A.
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Element
Testing/Prototyping
Outreach
Final Build
Travel to Competition
Total
Est. Cost
$13,972.23
$3,669.77
$4,202.33
$8,200
$29,922.13
Table 64 Preliminary Budget Summary
Funding Plan
A significant portion of the funding necessary for this project derives from a wide
range of University organizations and other community support functions. Thus
far, $11,500 in donations from the Tarleton President’s Circle, the Provost’s Office,
the Dean of the College of Science, and the Tarleton Foundation fund the
project. The Office of Student Research has provisions for the project amounting
to $17,000. USLI Science Mission Directorate (SMD) funding also stems from
NASA in the amount of $2,780. The total allocation for the project currently
amounts to $31,280.
Figure 70 - Initial Funding
Timeline
The Tarleton Aeronautical Team understands that a project of this magnitude
requires a great deal of time and dedication. The following schedule to meet the
requirements of the project serves as evidence in Figure 71. Gantt Charts detailing
the project timeline follow.
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Gantt Timeline
The following chart gives a visual representation of major project deliverables.
Figure 71 Project Timeline
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Testing Gantt Timeline
The Gantt chart below includes preliminary testing dates. Time is in the schedule
to allow for lab prototyping and testing. The team plans to conduct multiple test
launches including low altitude as well as high altitude test launches when
possible. Using data gathered from these test launches, the team performs a
failure analysis after each launch. These analyses are useful to optimize
successive launches and overall design.
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Figure 72 Testing Timeline
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Outreach Gantt Timeline
The chart below delineates the dates the team plans for educational outreach. The
star parties each involve a simple vehicle demonstration and a presentation about
the basics of rocketry. During class trips, team members travel to area middle
schools to actively engage students in safe, basic rocketry. The Tarleton Science
Olympiad consists of area middle school and high school students convening at
Tarleton to compete in science related activities. During the Science Olympiad the
team demonstrates a static motor test in addition to giving presentations explaining
basic rocketry. For more comprehensive information on the educational outreach
component of the project, please refer to the following outreach section of this
document.
Figure 73 Outreach Timeline
Tarleton Aeronautical Team 148
2012 - 2013 USLI Preliminary Design Review
Outreach Plan
Vehicle design, creation, and implementation are important components of this
competition. Conjunctively, the educational engagement portion of this project is
crucial, as its main goal is to promote enthusiasm for the necessary subjects that
relate to rocketry and other important STEM fields. The team’s plan is to host
several events for diverse audiences. All events aim to promote the global
necessity of math, science, engineering, and technology. Furthermore, the team
includes a vehicle launch with each event to provide a real world experience to
reinforce the addressed STEM concepts in the lesson portion of each event. The
team aims to encourage interest in the relevant subjects with the intent of
increasing the number of people that choose to pursue STEM related careers.
Figure 74 - Acton Middle School
Educational Outreach
Educational outreach targets three main audiences through a variety of events;
students, teachers, and the community as a whole. The team is currently
establishing contacts and scheduling dates to visit the local middle schools.
Several schools plan to participate in the educational outreach events already.
Outreach to students in the schools occurs through a classroom lesson or an
assembly style presentation. During the classroom sessions, small groups from
the team present an original interdisciplinary lesson over rocketry with an
emphasis on math and science.
The goal of these lessons is to demonstrate to students the importance of the
STEM subjects and their role in a variety of topics such as engineering and
rocketry. The necessity of these careers with companies such as NASA is a
primary focus. By working in a setting which allows for a smaller student to
presenter ratio, students are receive an opportunity to work closely with the team
members on a lesson which reinforces concepts learned previously in a novel
manner.
Tarleton Aeronautical Team 149
2012 - 2013 USLI Preliminary Design Review
Beyond the classroom lesson, the assembly format allows the team to
communicate the same information to students on a larger scale. This portion of
the outreach began at the request of some of the local middle schools. The
assemblies take place toward the end of the school day and involve multiple
classes and grade levels.
To conclude each school presentation, students join our team outside for a vehicle
launch. The rocket launch adds to the lesson by giving the students a visualization
of what they just learned. This increases students’ retention. In a classroom
setting, interactive, hands-on lessons encourage learning.
Educator Outreach
By aligning the lessons created for the classroom presentations with state and
national curriculum, these repeatable lessons remain relevant for reuse. The team
is working to create a live webcast of a vehicle launch for teachers to access in
their classes. This webcast allows teachers to use this online content as a real-life
application in their classroom. Furthermore, the team is communicating with
teachers throughout the state to distribute lesson plans. We hope to raise interest
in STEM fields by having teachers join our group. Lesson plans are available from
the Tarleton Aeronautical Team discussing rocketry and the importance of NASA.
Community Outreach
Star Party
Outreach beyond schools allows for students, teachers, and community members
to join in learning about rocketry and STEM concepts. The team coordinates with
Tarleton State University to co-host their Star Party event which occurs in both the
fall and spring semesters. The event includes a discussion about the program,
rocketry, the need for growth in the STEM fields, and a vehicle launch. The Star
Party is an open invitation event; the team reaches audiences with a range from
children to adults from local and surrounding areas.
Tarleton Regional Science Olympiad
The team will be at the eighth annual Tarleton Regional Science Olympiad on
February 23, 2013. Students participating in this event along with their sponsors
and family join the team for several vehicle launches including a static launch
demonstration. A presentation and question and answer session follow. The day
concludes with an awards ceremony. Again, the focus of this presentation is to
promote the STEM fields and reiterate their importance pertaining to the nation’s
progress.
Tarleton Aeronautical Team 150
2012 - 2013 USLI Preliminary Design Review
Participation Goal
The team expects to involve approximately 2,500 students in total. Comprehensive
feedback from teachers and students will be gathered through surveys. This
feedback helps the team alter presentations to ensure the quality of each event.
Outreach efforts for boy scouts, girl scouts, and after school programs are being
developed. The team members have a passion for the STEM fields, thus outreach
is an important goal.
Accomplished Educational Outreach
On October 5, 2012 members of the Tarleton Aeronautical Team traveled to
Granbury. Students at Acton Middle School and Granbury Middle School
participated in basic rocketry presentations. The team’s presentation explained
STEM fields and related careers. As part of Career Day at Acton Middle School,
the team specifically discussed careers available at NASA.
Bert led presentations featuring 7 Minutes of Terror, a NASA video highlighting
interviews with NASA engineers on the Curiosity Rover project. While gaining
exposure to career options with NASA, the students also learned the importance of
safety protocol. To emphasize the message of the video, the students were given
the opportunity to experience rocketry in a safe environment. Before the team
launched their rocket, each class was given the opportunity to launch two-liter
water bottle rockets.
The surveys reflected that the majority of the students were delighted with their
experience. In the survey, the students were asked to report whether they felt a
greater interest in Science, Mathematics, Engineering, or Technology, three things
they learned from the presentation, and what their favorite parts of the
presentation were. The data is given in Table 65 and illustrated in Figure 75.
Subject Interest
Science
Math
Engineering
Technology
Rocketry
None
Count
39
35
36
33
14
15
Table 65 Accomplished Educational Outreach
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2012 - 2013 USLI Preliminary Design Review
Subject Interest
Science
Math
Engineering
Technology
9%
8%
Rocketry
None
23%
19%
20%
21%
Figure 75 Subject Interest
Areas of learning include force concepts, propulsion, rocket construction, rocket
design, launch procedures, failure modes, qualifications for building rockets,
careers in rocketry, competitions in amateur rocketry, and information about
NASA. The categories with the greatest percentage of student learning were
rocket design, qualifications for building rockets, launch procedures, and
propulsion. This indicates that more time should be spent in future presentations
on the other learning categories, but further sampling is required. The data is given
in Table 66 and illustrated in Figure 76.
Presentation Learning Outcomes
Count
Forces
5
Propulsion
22
Construction
8
Design
31
Procedures
21
Failure Modes
13
Qualifications
24
Careers
14
Competitions
5
NASA
11
None
8
Table 66 Presentation Learning Outcomes
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2012 - 2013 USLI Preliminary Design Review
Presentation Learning Outcomes
Forces
Propulsion
Construction
Design
Procedures
Failure Modes
Qualifications
Careers
Competitions
NASA
None
7%
5% 3%
3%
13%
5%
9%
19%
15%
8%
13%
Figure 76 Presentation Learning Outcomes
When asked what their favorite part of the presentation was, the greatest number
of students responded in favor of the water bottle rocket activity. The data is given
in Table 67 and illustrated in Figure 77.
Favorite Part
Count
Launch
11
Launch Failure
3
Flight
5
Outside
10
Water Rockets
27
Big Rocket
21
None
7
Table 67 Favorite Part
Tarleton Aeronautical Team 153
2012 - 2013 USLI Preliminary Design Review
Favorite Part
Launch
Launch Failure
Flight
Outside
Water Rockets
Big Rocket
None
8%
13%
4%
25%
6%
12%
32%
Figure 77 Favorite Part
The surveys conducted at the two middle schools on October 5, 2012 were free
response. The Granbury events were the first conducted by the team. They
provided a wide variety of student responses concerning the presentation and
demonstration. This feedback will ultimately be used to formulate a comprehensive
and unbiased multiple-choice survey to be conducted at subsequent events. This
will boost the quality of questions posed at future presentations.
Conclusion
The team is eager to continue the design process and see the final product come
to fruition. Up to this point, the project is both challenging and rewarding. Although
this is the first time to compete in NASA’s USLI, the Tarleton Aeronautical Team is
confident in its ability to design a vehicle and payload to the customer prescribed
specifications. Creative, legitimate solutions have been posed and implemented,
defining the performance characteristics of the vehicle and payload systems as a
whole.
The vehicle design features reflect standard practices in amateur rocketry. All
components of the vehicle aim to achieve the goal of delivering the SMD payload
to one mile above ground level.
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2012 - 2013 USLI Preliminary Design Review
The SMD payload criterion calls for an atmospheric data gathering instrument that
meets all requirements as stated in the SOW. Controlling the sensors to make
measurements is among a long list of complicated problems creating a fully
functional payload. Other taxing problems include the orientation of the image
camera and the placement of the UV and solar irradiance sensors. The selfleveling camera design aims to take proper images by correcting offsets in the
payload orientation during descent. Pyranometers and UV sensors sit in an
opposing manner to effectively increase the field of view for solar irradiance and
ultraviolet measurement. Optical properties of the clear acrylic payload housing
must be addressed, and sensor performance testing must occur to verify this
housing structure selection. The payload will ultimately be inserted in a PCB due to
advantages in size, efficiency, repeatability and overall fidelity in design.
The educational outreach portion of the project is going extremely well. The
minimum requirement for the number of students to be reached was exceeded on
first day of this competition. The team continues to go above and beyond this
minimum requirement. The intention of the team is to expose as many students to
the STEM fields as possible.
Building upon the momentum spawned by the success in the 2012 CanSat
competition, along with support from the community and University, the team is
eager to progress through the design life cycle. Ultimately, flying the final vehicle
and payload on launch day will illustrate the team’s achievements. It is a true
testament to the abilities and ingenuity of the team members, and provides an
invaluable exercise in creating real-world engineering experience.
Tarleton Aeronautical Team 155
2012 - 2013 USLI Preliminary Design Review
Appendix A - Itemized Subsystem Budget
A.1 Structure System
Body
Part Number
Seller
Price Per
Unit
# of Units
Total
FNC5.5EL
WildMan
Rocketry
$47.50
1
$47.50
G12-5.5-60
WildMan
Rocketry
$38.84
6
$233.04
ACRCAT5.500ODX.250
ePlastics
$39.05
2
$78.10
500SHT0.125X48X96
ePlastics
$208.00
0.2
$41.60
G12-3.0-48
WildMan
Rocketry
$71.06
1
$71.06
PVCGRAY2.00LAM12x24
ePlastics
$147.82
0.25
$36.96
FBP5.5
WildMan
Rocketry
$7.60
5
$38.00
Couplers
G12CT-5.5
WildMan
Rocketry
$47.03
2
$94.06
Centering Rings
FCR5.5-3.0
WildMan
Rocketry
$8.55
3
$25.65
Epoxy
4500Q
WildMan
Rocketry
$69.00
0.25
$17.25
Motor Retainer
RA75
Wildman
Rocketry
$52.00
1
$52.00
$15.00
2
$30.00
Grand
Total
$765.22
Nose Cone
Fiberglass Body
Tube
Acrylic Body Tube
Sheet for Fins
Motor Tube
Bulk Plate
(Payload)
Bulk Plate Standard
Shipping
A.2 Recovery System Budget
Tarleton Aeronautical Team 156
2012 - 2013 USLI Preliminary Design Review
Proposed Selection
Distributor
Item Number
Unit
Cost
Proposed
Quantity
Typical
Unit Cost
Main Altimeters
Featherweight
Altimeters
Raven3
$155.00
2
$155.00
Backup Altimeters
PerfectFlite
StratoLogger
$79.95
3
$79.95
Electric Matches
Coast Rocketry
Daveyfire N28BR
$2.95
4
$11.80
FFFFg Black Powder
Goex
Goex 4F Black
Powder
$15.75
1
$15.75
Black Powder Ejection
Charge Holders
Aerocon Systems
BPSmall
$3.00
1
$3.00
Swivels
Commonwealth
Rocketry
SWLDK80
$1.99
2
$3.98
Main Shock Cord
Giant Leap
Rocketry
Tubular Kevlar
$37.99
1
$37.99
Drogue Shock Cord
Giant Leap
Rocketry
Tubular Kevlar
$31.49
1
$31.49
Main Parachute
Recovery
Technologies
Premium with Spill
Hole
$145.00
1
$145.00
Flameproof Main
Parachute
Deployment Bag
Rocketman
Enterprises
DB8
$40.00
1
$40.00
Drogue Parachute
Recovery
Technologies
Premium with
Spillhole
$16.00
1
$16.00
Flame-Proof Drogue
Parachute
Deployment Bag
Rocketman
Enterprises
DB2
$25.00
1
$25.00
U-Bolts
Sunward
Aerospace
U-Bolt Assembly 0.25in. (compact)
$4.29
2
$8.58
Quick Links
Commonwealth
Rocketry
0.25in. Stainless
Steel Delta Quick
Link
$2.99
4
$11.96
Shear Pins
Missile Works
2-56 Nylon Shear-Pin
(10 pack)
$1.00
1
$1.00
Arming Switches
Featherweight
Altimeters
Featherweight
Magnetic Switch
$25.00
5
$125.00
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2012 - 2013 USLI Preliminary Design Review
Hand-held Rotary Fan
Anemometer
WeatherShack
SpeedTech WM-300
$154.95
1
$154.95
GPS
Big Red Bee
BeeLine GPSPackage Deal
$289.00
2
$578.00
Total
$1,444.45
A.3 Payload Budget
Product
Arduino 2560-R3
Price
Quantity
Total
$58.95
1
$58.95
micro SDHC
$9.99
2
$19.98
Adafruit 254
$15.00
2
$30.00
BMP180
$15.00
1
$15.00
MS5611-01BA03
$29.99
1
$29.99
HIH4030
$16.95
1
$16.95
HH10D
$9.95
1
$9.95
SP-110
$169.00
2
$338.00
TSL2561
$12.50
2
$25.00
SU-100
$159.00
2
$318.00
TOCON_ABC3
$148.00
2
$296.00
HS-85BB+ Mighty Micro
$19.99
2
$39.98
Adafruit 397
$42.00
1
$42.00
VCC-003-MUVI-BLK
$49.95
1
$49.95
Xbee-PRO XSC S3B
$42.00
1
$42.00
LS20031
$60.00
1
$60.00
LCD-11062
$34.95
1
$34.95
ADXL345
$27.95
1
$27.95
$6.65
8
$53.20
Ultralife U9VLBP
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2012 - 2013 USLI Preliminary Design Review
Mouser 534-3427
$10.22
2
$20.44
Adept A1E
$29.95
1
$29.95
$1,558.24
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2012 - 2013 USLI Preliminary Design Review