Student Launch 2013-2014
Critical Design Review
Project Advance
2/28/14
SUMMARY OF CDR REPORT
Team Summary
Team Name: Project Advance
School:
University of Central Florida
Address:
Department of Mechanical, Materials and Aerospace Engineering
4000 Central Florida Blvd,
Orlando FL, 32816
Team Mentor: Anthony Laiuppa, NAR 9399, Level 2 Certified
Launch Vehicle Summary
Size and Mass: 91 in length, 3.15 in diameter, 8.5 lb.
Motor: Cesaroni K-660 Sparky
Recovery System:
The rocket will feature a dual deployment recovery system powered by a PerfectFlite
Stratologger. The recovery will be composed of two parachutes. One being an X-form drogue
chute to be deployed at apogee, and the other to be an 84 inch semi hemispherical main
parachute to be deployed at 750 feet. Using these parachutes and projected altitudes for
deployment the rocket will descend at a rate of 17.24 feet per second according to our rate of
descent calculations. This falls well below the NAR minimum rate of descent which is at about
20 feet per second. In order to assure the chutes deploy the altimeter will be wired to ejection
canisters on each end of the electronics bay containing approximately 1.5 ounces of black power.
Based on previous experience and the power of black powder this should be more than enough to
deploy the recovery. Additional details can be found further down under the Recovery System
heading.
Rail size - The vehicle will be launched on an eight foot long rail that is one inch in diameter.
CDR Milestone Review Flysheet
Payload Summary
3.1 The rocket will include a camera that will scan the ground for landing hazards and the data
will be transmitted in real time to a ground station and analyzed using custom made on-board
software.
Video will be taken using a 5MP camera module. The on-board software will be
processed using a Raspberry Pi and will, in real time, analyze frames in the video for any hazards
and video captures along with data that will be transmitted to a screen on a ground station. A 3G
Dongle receiver will be connected to the Raspberry Pi video output and transmit video data to
the ground station with any data about ground hazards overlaid onto the video.
3.2.1.3 This project will examine the airframe, propulsion, and electrical systems through
structural and dynamic analysis during boost.
Strain gauges will be placed on the body tube to measure strain on the structure of the
rocket through voltage readings. The electrical system will be analyzed by connecting voltmeters
at various points along the circuit. To analyze propulsion, an accelerometer will be placed on the
body tube of the rocket to measure acceleration during burn and an altimeter will be fixed to the
rocket to measure its velocity. Taking the positive y-direction to be vertically straight up, the
sum of the forces on the rocket in the y-direction during burn is found to be
1
Eq. 1 ↑ ∑ 𝐹𝑦 = 𝑚𝑎 = 𝑇 − 𝑚𝑔 − 2 𝜌𝐶𝑑 𝑣 2 𝐴
Eq. 2
1
2
𝑚𝑎 + 𝑚𝑔 + 𝜌𝐶𝑑 𝑣 2 𝐴 = 𝑇
where “m” is the mass of the entire rocket, all its components, the motor and motor casing, and
half of the propellant, “a” is the acceleration of the launch vehicle measured by the
accelerometer, “v” is the velocity of the rocket measured by the altimeter, “g” is the acceleration
due to gravity, “” is the density of air, “A” is the cross-sectional area of the rocket, “” is the
coefficient of drag found using computational fluid dynamic simulation in Solidworks, and “T”
is the thrust generated by the motor. From equation 2 the thrust of the motor during burn can be
determined assuming a perfectly vertical launch with no wobble. The burn time of the motor will
be determined from the acceleration vs. time graph generated by the accelerometer. The time
from the point on the graph that shows acceleration to the point on the graph that shows the start
of deceleration will give the total burn time of the motor. The impulse generated by the motor
can then be found using equation 3
Eq. 3 tT=I
where “t” is the burn time of the motor, “T” is the thrust of the motor, and “I” is the impulse
generated by the motor.
3.2.2.2 Aerodynamic analysis will also be performed on protuberances from the structure which
will include computational fluid dynamic simulation and wind tunnel testing.
The team will examine the aerodynamics of the rocket by using computational fluid
dynamic simulation in SolidWorks, especially to determine the effect of protrusions from the
structure such as the video cameras.
Changes made since PDR
Changes made to vehicle criteria
The airframe has now been changed to solid fiberglass. This is to account for the
structural rigidity during supersonic flights. As it is a much stronger material than quantum
tubing it is less prone to cracking and flexing during flight. We have also changed the design of
the fins to a fin that is made of G10 fiberglass in a clipped delta shape. With a span of 4 inches
and a root chord of 6.5 inches our fin flutter calculator has given us a result that the fins will not
shred and hold up well. This is an improvement on the previous fins that were much more
unstable. We have also moved some of the electronics towards the top of the rocket as it adds
stability to the rocket and also allows us to fit our electronics in better given our space allotted.
Item Changed
PDR
CDR
Reason
Hazard
Detection
Equipment
Arduino Uno and Video
Experimenter Shield with
Radio Transmitter
Raspberry Pi
Model B and 3G
Dongle
Raspberry Pi was found to
be more reliable in
processing video data
AirFrame
Quantum Tubing
Fiberglass
Improved structural
rigidity during supersonic
flights.
Fins
Trapezoidal Fins
Clipped Delta G10 Fiberglass
Reducing fin flutter for
stable flight
Changes made to payload criteria
The cameras will now be tucked inside the nose of the launch vehicle and spring loaded
to deploy when the drogue chute is deployed. The reasoning for putting the cameras inside the
nose is to prevent them from possibly shearing off during launch. The nose cone will always be
pointing downward during the recovery processes of the launch so the camera will always be
facing the correct direction.
There will now be two electronics bays, one in the nose cone area, the other in the
original electronics bay location of the body tube. The justification for putting an electronics bay
in the nose has to do with the location of the cameras. These electronics will be located in the
nose, therefore the wiring and battery must be able to reach these electrons without being
damaged during the deployment of the recovery system. Team Advance discovered that the best
way to do so would be to simply make another electronics bay in the nose. The other electronics
bay will have to exist further down the body tube in order to gather data from a strain gauge that
will be placed on the structure of the launch vehicle. Both electronics bays will be shorter than
the original planned size from the PDR since they will both house less electronics, so the overall
length of the launch vehicle will not be affected.
Team Advance has also decided to use a Raspberry Pi instead of an Arduino to send
video footage to the ground station. Through testing on a personal Raspberry Pi the software
engineer of Project Advance has determined that it will be easier to program a Raspberry Pi to
send video footage to the ground station than it would be to program an Arduino to do the same.
However, an Arduino will still be used in the electronics bay of the body tube to record data
from the strain gauge and accelerometer.
Changes made to project plan
Subscale and full scale assembly and test launches could be delayed. Team Advance is
still waiting on parts to arrive.
III) Vehicle Criteria
Design and Verification of Launch Vehicle
#
Requirement
Feature
Verification
1.1
The vehicle shall deliver the
research payload to a
predetermined altitude of
8,500 ft. above ground level.
The rocket will have a payload The rocket will carry
bay capable of housing the
the payload during the
research system.
full scale test launch.
1.2
The vehicle shall carry one
commercially available,
The design of the rocket will
designate a space for an
The altimeter will be
verified during the half
barometric altimeter for
recording of the official
altitude used in the
competition scoring.
altimeter.
scale and full scale test
launch.
1.3
The launch vehicle shall be
The rocket, fins, and payload
designed to be recoverable and bay will be made of strong
reusable.
materials able to withstand the
launch and landing. The
rocket will decouple and
implement parachutes. The
cameras will allow the team to
monitor the location where the
rocket lands.
The decoupling system
and parachutes will be
inspected prior to
launch. The integrity of
the materials will be
tested during the half
scale and full scale test
launch.
1.4
The launch vehicle shall be
capable of being prepared for
flight at the launch site within
2 hours, from the time the
Federal Aviation
Administration flight waiver
opens.
The team will practice
preparing the rocket for flight.
Team preparation
efficiency will be tested
during the half scale
and full scale test
launches.
1.5
The launch vehicle shall be
capable of remaining in
launch-ready configuration at
the pad for a minimum of 1
hour without losing the
functionality of any critical
on-board component.
The vehicle will contain
sufficient batteries to supply
power to the research systems
and payload for over an hour.
The circuits involved in
the research system will
be analyzed for power
consumption.
1.6
The launch vehicle shall be
capable of being launched by
a standard 12 volt direct
current firing system.
The rocket will be designed to
meet this criteria.
This will be tested and
verified during the half
scale and full scale test
launches.
1.7
The launch vehicle shall
require no external circuitry or
special ground support
equipment to initiate launch
other than what is provided by
Range Services.
The rocket will be built to be
compatible with the
equipment provided by Range
Services.
Launch equipment
compatibility will be
verified during the full
scale test launch.
1.8
The launch vehicle shall use a The team will be using an
The motors
commercially available solid
NAR-certified Cesaroni K-660 specifications will be by
motor propulsion system using Sparky motor.
the RSO.
ammonium perchlorate
composite propellant as
approved by the National
Association of Rocketry.
1.9
Pressure vessels on the vehicle The pressure vessels will be
shall be approved by the RSO designed to the specified
and shall meet criteria 1.9.1 to standards.
1.9.4.
1.10 All teams shall successfully
launch and recover their full
scale rocket prior to FRR in its
final flight configuration and
criteria 1.10.1 to 1.10.5 must
be met in the full scale
demonstration flight.
This will be verified
during the full scale
launch.
The rocket will be flown with
the payload and full scale
motor and will not be altered
following the full scale
demonstration flight.
This will be verified
during the full scale test
launch.
2.1
The launch vehicle shall stage
the deployment of its recovery
devices, where a drogue
parachute is deployed at
apogee and a main parachute
is deployed at a much lower
altitude.
The team will design the
rocket to deploy these systems
at the proper altitudes by
activating altimeters at apogee
for the drogue and 1000 feet
for the main parachute.
These systems will be
tested during the half
scale and full scale test
launches.
2.2
The parachute system(s) shall
be designed and manufactured
by the team.
The team will design the
system by combining research
and prior experience.
Mentor Anthony
Laiuppa will give a
written statement
verifying that the
parachute system has
been designed and
manufactured by the
members of Project
Advance.
2.3
At landing, each independent
sections of the launch vehicle
shall have a maximum kinetic
energy of 75 ft-lbf.
The parachutes will be
designed to minimize kinetic
energy by the time of impact.
The parachutes will be
analyzed in programs
like Rocksim and tested
during half scale and
full scale test launches.
2.4
The recovery system electrical
circuits shall be completely
independent of any payload
electrical circuit.
The recovery system electrical
systems will be properly
isolated from the payload
electrical system.
The circuitry will be
tested on the ground
before and after the half
scale launch as well
before the full scale test
launch.
2.5
The recovery system shall
contain redundant,
commercially available
altimeters
The drogue and parachute will
have redundant sets of
altimeters.
The altimeters will be
tested on the ground
prior to launch using a
vacuum chamber.
2.6
Each altimeter shall be armed
by a dedicated arming switch
on the rocket airframe.
A screw switch will be
designated for each altimeter.
The circuitry will be
tested on the ground to
ensure that each
altimeter is connected to
a single screw switch.
2.7
Each altimeter shall have a
dedicated power supply.
A battery will be designated
for each altimeter.
The circuitry will be
tested on the ground to
ensure that each
altimeter is receiving
power from a separate
battery.
2.8
Each arming switch shall be
capable of being locked in the
ON position for launch.
Switches are designed to lock
securely during all phases of
flight.
The switches will be
tested during full and
half scale test launches.
2.9
Removable shear pins shall be
used for both the main
parachute compartment and
the drogue parachute
compartment.
The main parachute and
drogue will use removable
shear pins that will be broken
by black powder charges.
The shear pins will be
tested on the ground as
well as during half scale
and full scale test
launches.
2.10 An electronic tracking device
shall be installed in the launch
vehicle and on any untethered
rocket section or payload
component and it shall
transmit the position of the
tethered vehicle or any
independent section to a
ground receiver.
GPS will be installed on the
components of the vehicle.
The accuracy of the
location transmission
will be tested during
half scale and full scale
test launches.
2.11 The recovery system
electronics shall not be
adversely affected by any
other on-board electronic
devices from launch until
landing.
The recovery electronics
system will be properly
isolated from other on-board
electronic devices within the
rocket.
The circuitry will be
tested on the ground
before and after the half
scale launch as well
before the full scale test
launch.
3.1
The rocket will include a
camera that will scan the
ground for landing hazards
and the data will be
transmitted in real time to a
ground station and analyzed
using custom made on-board
software.
A camera connected to a
Raspberry Pi with image
processing software will send
hazard data using a
transmitter.
The camera will be
tested on the ground as
well as during half scale
and full scale test
launches.
3.2
The payload shall be designed
to be recoverable and
reusable.
The payload will be encased
in a robust material such as
metal in order to prevent
damage upon landing.
The durability of the
payload housing will be
tested during the half
scale and full scale test
launch.
3.3
The payload shall imcorporate
requirements 3.2.1.3 and
3.2.2.2 in addition to
requirement 3.1.
The payload will incorporate
all three systems as described
in detail in the Payload
Criteria section below.
All payload systems
will be tested
throughout the design,
build and test launches.
3.4
The team that demonstrates
the highest level of fidelity in
meeting requirements 3.4.1 to
3.4.10 will have the chance to
fly a microgravity payload on
the reduced gravity aircraft.
Not applicable.
Not applicable.
4.1
Each team shall use a launch
and safety checklist.
The team will work with the
safety officer to create these
checklists.
Launch and safety
checklists will be
verified by the safety
officer and used during
all launches.
4.2
Students on the team shall do
100% of the project, including
design, construction, written
reports, presentations, and
flight preparation with the
exception of a few minor
activities.
Project Advance will complete Team mentor Anthony
all aspects of the project.
Laiuppa will give a
written statement
verifying that Project
Advance completed all
aspects of the project.
4.3
The team shall provide and
maintain a project plan to
minimally include project
milestones, budget and
community support,
checklists, personnel assigned,
The team will work together
to plan these tasks for
completion in a thorough and
timely manner.
The plan will be used
throughout the entirety
of the project in order to
stay on task.
educational engagement
events, and risks and
mitigations.
4.4
Each team shall identify a
“mentor” prior to the PDR
who is certified by the
National Association of
Rocketry, owns the rocket for
liability reasons, and must
travel with the team to the
launch.
Team mentor Anthony
Laiuppa, who is certified by
the National Association of
Rocketry, will have ownership
of the rocket and will be
accompanying the team to the
launch.
The team mentor will
verify that these
requirements are met
and travel with the team
to the launch.
4.5
All team members attending
launch week activities shall be
identified by the Critical
Design Review.
Team members will confirm
their ability to travel and
attend launch week activities
before CDR.
The team will verify the
final list of members by
the time of CDR.
4.6
Foreign National team
members shall be identified by
the Preliminary Design
Review and may be separated
from their team during launch
week activities.
Project Advance will verify
citizenship of all team
members and explain possible
circumstances.
The citizenship of all
team members will be
verified for launch week
activities.
4.7
During test flights, teams shall
abide by the rules and
guidance of the local rocketry
club’s RSO.
Team members will follow
any rules set by the local
rocketry club’s RSO and
follow any guidance given.
Any rules set by the
RSO will be reviewed
by all members.
4.8
The team shall engage a
minimum of 200 participants
(at least 100 of those shall be
middle school students or
educators) in educational,
hands-on science, technology,
engineering, and mathematics
(STEM) activities.
Project Advance will engage
fifth grade and middle school
students in two events that
will allow the students to learn
about STEM fields through
interactive presentations and
hands on activities.
These events will be
documented and
verified through photos,
videos and the
Educational
Engagement Form.
4.9
The team shall develop and
host a Web site for project
documentation.
The team will maintain its
website and any
documentation will be posted
to it.
Team members will
verify that each
document is posted to
the website once a final
draft is created.
Table 1
Flight Reliability and Confidence
Mission statement, requirements, and mission success criteria
The mission of Project Advance is to send a rocket to an altitude of 8,000 ± 500 feet and
return the vehicle to the ground in a manner such that it can still be used in the future. Upon
descent the rocket will deploy onboard cameras that will scan the ground for landing hazards and
the data will be transmitted to a ground station in real time in order for data to be analyzed. This
project will also perform an examination of the airframe, propulsion and electrical system
through structural and dynamic analysis, as well as aerodynamic analysis on protuberances from
the launch vehicle. The mission of the project is also to serve at least 200 individuals in the local
community through STEM related outreach such as workshops and events. Mission success is all
dependant on the rocket’s ability to launch to 8,000 ± 500 feet in altitude and safely return to the
ground with the capability of being launched again without any major repairs to the vehicles.
Mission success is also dependent on the ability of the onboard cameras to deploy properly and
clearly send information about possible incoming landing hazards to a ground station in real
time. It is also required that the sensors such as the strain gauges, accelerometer, and barometric
pressure sensor function properly in order to extrapolate data from them. This means the
raspberry pi must collect data from the strain gauge while the arduino collects data from the
other strain gauge and the accelerometer. Additionally the barometric pressure sensor will collect
the data of its own accord and store it within its internal memory. Electrical ground tests using
voltmeters and CFD analysis through SolidWorks must be performed to show that the launch
vehicle can safely operate with no concerns. Additional testing concerning the ability of the
programmed devices to run properly given constraints defined by the programmer will also be
taken into account. Finally as a last measure of success, Project Advance must reach out to at
least 200 individuals in the community through STEM outreach events, with at least 100 of them
being middle school students or educators.
Major milestone schedule
Final Design Completed - February 28, 2014
Subscale Assembly Completed - March 7, 2014
Subscale Test Launch - March 8, 2014
Full Scale Assembly Completed - April 11, 2014
Full Scale Test Launch - April 12, 2014
Full Scale Alterations Completed - April 30, 2014
Rocket Ready for Shipment - May 3, 2014
Rocket Shipment - May 4, 2014
Unpack Rocket - May 14, 2014
Rocket Inspection - May 15, 2014
Launch Setup - May 16, 2014
Launch and Recovery - May 17, 2014
Post Launch Inspection - May 18, 2014
Design at a system level
Final drawings and specifications (detailed drawings)
Final analysis and model results, anchored to test data
Unfortunately, parts for the subscale and full scale launch vehicles have not yet
arrived. Due to this set back only ground tests have been performed up until this point
using parts and equipment that was already available to the team. The extent of the
ground tests that have been performed are described below in the section heading “Test
description and results.” Final model results anchored to test data have been found
digitally through the use of digital means. The three digital tests to be elaborated upon
below have given us the following three conclusions. The first being that our rocket is
stable, the second being that our finals will not flutter enough to induce shredding and the
third that our airframe will not flex in a manner that will cause destruction.
Test description and results
Black powder ejection charge testing was performed on a full scale test rocket
originally built for another competition. The rocket’s diameter and length matched that of
the Pioneer and was modified to match the mass of the Pioneer. This was done because
parts for the full scale rocket have still not yet arrived. The test was performed with two
charges, one to eject the drogue chute, the other to eject the main chute. This was done to
see how much black powder was needed to confidently eject the recovery system. It was
found that one, 1.5 g. black powder charge each was sufficient enough to deploy the
recovery system. Although 1.5 g of black powder was enough, tests were also performed
with 2 g of black powder to include a factor of safety into the design. The test also
showed that two, 2 g black powder charges are not too much force that would possibly
damage the vehicle, electronics bay, or components of the recovery system such as the
shock cord and parachutes. With these results the team has decided to use two, 2 g black
powder charges for the full scale rocket.
As far as digital tests go the first of the three is the stability testing, it was
modeled based on data provided by RockSim. RockSim tested the center of pressure and
center of gravity using advanced Barrowman equations that were developed by James
Barrowman a former NASA engineer. The results led us to a static stability margin of
2.67. This is only marginally over stable. This is however a good static stability margin as
a static margin of two body diameter lengths between the CG and CP is the industry
accepted safety margin used by NAR and Tripoli for high power rocket design.
Additionally the Barrowman equations model a worst case scenario so in the event that
our rocket was minorly unstable there is still a chance that it will still be stable. So as a
result of the modeling the data does reinforce the thought that the rocket is stable as it
meets the above criteria.
The next modeling test ran to be backed up with data was that of fin flutter. Upon
reviewing the old design and then following it with research we were able to model that
we needed to revise the design. We did this by reading NACA Tec Note 4197 and “The
Flutter Problem, Modeling Unstable Aeroelastic Vibrations in Rocket Fins” a research
paper by Vincent San Miguel a Penn State student. Using the equations and knowledge
garned in these papers we were able to combine that collected information with a
spreadsheet calculator to produce data that would model and project fin flutter in our fin
design. We found that by shortening the span, increasing the root cord and changing the
shape, that flutter would cease to be a problem. This test has yielded the results, that are
based in data, that by using a clipped Delta Shape, that fin flutter will not be a problem in
this situation. Infact it was found that we had a 20% margin of safety giving us a 120%
stability when faced with fin flutter.
Another modeling simulator we ran was that of Flex- Roc. It models and projects
the aeroelastic divergence of the body tube by using a program that runs iterations of a
complex equation to give you the degree of safety concerned with aeroelastic divergence.
Through this we input the numbers correlating to our rockets dimensions and materials to
test for aeroelastic divergence in our rocket during its flight. The results of the test were
backed in data that showed a trend that the rocket would not be prone to flexing or
contortion during flight. This was shown by data pointing towards a safety margin of
20% giving our rocket a 120% total structural rigidity. This data and test has backed up
the decision to use a fiberglass airframe in the final design by confirming the idea that
since fiberglass is so rigid it is not prone to flexing.
Tests performed on the electronics bay within the nose cone include the following:

Running Python scripts to check if Internet access is available. If so, the Raspberry Pi will
connect to the Internet. The scripts will continuously running while in flight to check if the
Internet can be reached so that video can be relayed to the ground station.

Testing cameras to ensure that hazard detection software performs correctly. Considerations will
be made to assure that the camera and software work together. Hazard detection tests include
accuracy and precision in detecting landing hazards.
Test results for nose cone electronics bay:

Python scripts successfully check for Internet access and connect the Raspberry Pi to the
Internet if it is detected.

Hazard detection software and hardware were found to work correctly with each other.
Some errors were found in hazard detection software where false positives were detected,
but overall hazards were mostly detected accurately and precisely.
Final motor selection
The motor of choice for Project Advance is the Cessaroni K-660 Sparky. This motor was
selected based off the results of previous rocket projects performed by members of Team
Advance including the SEDS 10k competition and personal projects of team members. The team
is confident that the Cessaroni K-660 will launch the vehicle to the target altitude of 8,500 ft.
based off results of these past projects. Also, Rocksim rocket simulator was used to simulate
launches and estimate the altitude that would be reached with the Cessaroni K-660 motor. The
rocket’s size, mass, and design were put into Rocksim and run. The results showed that the
rocket would easily make the desired 8,500 ft. altitude.
In Rocksim the simulations showed that the rocket would overshoot the 8,500 ft. altitude,
however, the team has much experience with Rocksim and has found that the simulations will
often predict that the launch vehicle will launch to a higher altitude than it actually will. From
previous experience with Rocksim the team is confident that the Cessaroni K-660 Sparky will
not overshoot the target altitude, but will provide the required thrust to launch the rocket close
8,500 ft.
Approach to workmanship as it relates to mission success.
The teams approach to workmanship will be following part of the UCF Creed of
Integrity, Community, Creativity, Scholarship and Excellence. In this case Team Advance will
be specifically focusing on the word Excellence. As a team we would like to excel in
workmanship as we represent UCF through our attempt at mission success. It is not only the best
way to represent UCF by following the UCF creed, but also the best way to chase after mission
success through workmanship by having excellence in all the team does. This means working
hard to insure that the rocket isn't just adequate to achieve the mission, but that it excels in all it
does. It is vital that all tasks are completed using a high level of precision as to avoid making
mistakes that would compromise the functionality or the scientific objectives of the mission.
Planned additional component, functional, or static testing.
As soon as parts arrive the team will immediately begin construction of the subscale rocket
and prepare the electronics bay for a subscale test flight. The purpose of the subscale launch will
be to test the accelerometer, strain gauges, cameras, altimeters, recovery system, Raspberry Pi,
and Arduino to see if these components function properly in flight and that proper data is being
gathered and stored for later analysis. The team will also perform all the calculations and
analysis from the subscale launch just as they will for the competition to ensure everything goes
smoothly on launch day. This includes using the acceleration data from the accelerometer and
velocity gathered from the altimeter to find the thrust and impulse generated by the motor. Strain
gauge data will also be converted from voltage measurements into strain measurements to see
how much strain the structure of the launch vehicle endures during launch and if reinforcements
to the vehicle structure need to be made for the full scale rocket. The subscale launch will also
show whether or not improvements need to be made to the design of the recovery system, such
as insuring that the parachutes are safe from burning and that they deploy properly and at the
proper altitudes.
Status and plans of remaining manufacturing and assembly.
All assembly of both the subscale and full scale rockets still need to be completed due to the
late arrival of parts. As soon as parts arrive Team Advance will begin quick construction of the
subscale and perform a test launch as soon as possible. Based on the results of the subscale
necessary adjustments will be made and full scale assembly will begin. The team will strive to
build both the subscale and full scale as quick as possible without sacrificing safety or quality.
Design Integrity
Suitability of shape and fin style for mission
Previously the issue of fin shape and size was mentioned during the preliminary design
presentation of the rocket. Upon researching the issue it was found that the fins were unstable as
they were trapezoidal and had to long of a span. In order to properly research the issue of fin
flutter a published paper from Penn State University by Vincent San Miguel titled “The Flutter
Problem, Modeling Unstable Aeroelastic Vibrations in Rocket Fins”, as well as NACA Tec Note
4197 were used along with the accompanying formulas and spreadsheets set up to model our fin
flutter for analysis. Upon using these equations to model various forms of fins, two conclusions
were reached. The first is that the fin flutter equations provided are not accurate for supersonic
flight and thus are not feasible for purposes of the project as they would require a fourth order
partial differential equation to properly model. Second was that our current shape and
dimensions was not suitable for the rocket even reaching speeds up to mach before the equations
begin to become inaccurate. In order to amend this we have decided to retain the material being
used which is G10 fiberglass and plenty strong however, we have changed the shape to that of a
Clipped Delta. With a smaller span and slightly elongated root chord it ran through the fin flutter
sheet with numbers showing our fin would be stable at subsonic speeds and shall not flutter
excessively or shred. Infact it was more than stable as the calculator produced a 120% stability
rating which coincides with current research belonging to the calculator stating you want at least
that 20% margin to ensure the rigidity of the fins. Additionally the Clipped Delta was picked as
upon impact the tip of a traditional Delta fin would be the point of weakness and be especially
prone to breaking. In order to combat this a Clipped Delta was picked as it reduces the change of
this breaking upon impact. Thus we have found the suitable shape and fin style for the mission to
be a Clipped Delta constructed from G10 Fiberglass. The following diagrams are related to
structural integrity and fin flutter. (Image courtesy of Contest - Roc Yahoo group.)
Image showing Lack of Divergence after computation
A graph of a non-divergent rocket based upon the data. Increase it by an order of E01 and the
results are accurate towards out rocket.
Fin Flutter Calculator Courtesy of Duncan McDonald
Equation Courtesy of Vincent San Miguel
The result stating the G-10 will not shred.
Final chart showing the integrity of our rocket
Fins, bulkheads, and structural elements and proper assembly procedures, proper
attachment and alignment of elements, solid connection points, and load paths
To begin with the material of the fins will be G10 fiberglass as it the second strongest
commercial grade material the team has access to. It has been chosen as our calculations and past
knowledge shows it to be rigid enough to withstand the duress and strains of high speed flights.
Next the team has decided to chose carbon fiber as the material of choice for bulkheads. This is
largely due to the availability and capability of the team to effectively manufacture the bulkheads
in house. Additionally as some bulkheads will act as anchors for various pieces related to the
electronics bay or shock cord, carbon fiber is the optimal choice as drilling into any material to
anchor these items will weaken it fundamentally. However given the strength of carbon fiber it
should be able to withstand and tension from the shock cords pulling on the bolts anchored in
them, while also maintaining much more structural rigidity when drilled into than if balsa wood
or fiber glass had been chosen. Through the use of the Flex-Roc structural integrity calculator for
our rockets airframe, provided by Contest-Roc yahoo group, a reading was given stating that our
structure is not prone to aeroelastic divergence. It used a software running iterations based upon
numbers and densities correlating to our rocket and its components to estimate if aeroelasticity
divergence would occur to our body tubing during the flight. The calculator gave us the result
that the rocket had a safety factor of 120%. Meaning it is more than structurally safe for the
flight based on the projected velocity and calculations ran by the program. This reinforces the
thought that as fiberglass is inelastic it is not prone to bending or flexing. In order to further
insure the structural integrity of the airframe the body tubing will have at least 4 inches of
coupler connecting the parts. This lines up with common high power construction techniques.
We will also use another high power construction technique by putting our fins through the wall
to connect them directly to the motor mount.
Sufficient motor mounting and retention
The motor casing will be mounted inside a motor mount which will be secured into the
rocket by being epoxied in. After that it is as simple as installing an AeroPak 54mm motor
retainer to hold the motor casing inside the motor mount.
Status of verification
The status of the verification is that the rocket has been verified to be stable and
structurally sound. This has been determined by multiple calculators and programs modeling the
points of failure across the various subsystems of the rocket and analyzing probable events to
cause failure. Using this, the team was able to mitigate these issues by working around them in
various ways described throughout the paper.
Drawings of the launch vehicle, subsystems, and major components
Mass Statement and expected mass growth
The estimated mass of the final design of the rocket is to be 136.9992 ounces. The
NAR code defines the minimum acceptable thrust as 5:1. Thus by that standard and the formula
to determine this ratio we have concluded the rocket would have to be more than 30 pounds to
exceed this standard. This is all founded through RockSim which has given us accurate values
for all of the rockets components we do not possess to weigh such as the body tube. It has also
given us the ability to add in mass overrides based on the electronics systems we do possess
allowing us to get a more accurate reading. The expected mass growth between CDR and
delivery of the product is going to be between 3 and 4 pounds. This is just to account for any
additional pieces it may take to power electronics and additional weight that RockSim cant
predict like putty applied to the fins to fillet them.
Overall the vehicles mass is quite resilient to most changes in mass and even projecting 4
pounds of mass growth is a worst case scenario. So the rocket should possess a more than
adequate amount of thrust to get it off the ground and to the projected altitude.
Discuss the safety and failure analysis.
Failure analysis: The failure analysis begins with the outermost layers and then proceeds
inwards. The failure analysis for the airframe can be broken down into three parts. The point of
failure regarding the stability of the rocket, the point of failure regarding its fins, and the point of
failure that is the body tube itself.
The failure analysis that has been done on the stability of the rocket has been modeled in
RockSim and the data has been collected accordingly. Through RockSim we calculated the
center of pressure and center of gravity using the advanced Barrowman equation that was
developed by NASA engineer James Barrowman. Standard static stability margin of 2 body
diameters is the industry accepted safety margin used by NAR and Tripoli. As we are close to
this with a static stability margin of 2.67 the rocket should be fine. Static stability margin is a
massive part of the failure analysis as it is a failure of the rocket as a whole if it were to head off
in a wrong direction due to instability. It can only be mitigated by properly designing the rocket
to be stable which in this case, the rocket has met the above criteria.
The next point of failure of the rocket will be failure point involving the fins. As the
rocket is moving upwards if the fins are not of the proper shape as higher speeds are reaching the
fins will begin to flutter. In some catastrophic situations the fins can flutter violently enough to
shred and render them and the lower region of the rocket useless. In order to combat this we first
began to look at our fin design from the PDR. We took note of its shape and span. Upon
reviewing the old design and then following it with research we were able to conclude that we
needed to revise the design. We did this by reading NACA Tec Note 4197 and “The Flutter
Problem, Modeling Unstable Aeroelastic Vibrations in Rocket Fins” , a research paper by
Vincent San Miguel who is a Penn State student. Using the equations and knowledge garned in
these papers we were able to combine that with an Excel spreadsheet calculator to produce data
that would model and project fin flutter in our fin design. We found that by shortening the span,
increasing the root chord and changing the shape, that flutter would cease to be a problem. This
test has yielded the results, that are based in data, that by using a clipped Delta Shape, that fin
flutter will not be a problem in this situation. Infact it was found that we had a 20% margin of
safety giving us a 120% stability when faced with fin flutter. This in turn mitigates the point of
failure that is known as fin flutter during flight.
There is also a point of failure in the way the fins are attached to the rocket. Using a
slotted body tube the fins will be mounted to the motor mount of the rocket by a method referred
to simply as “through the wall fins.” The fins will then be filleted by an epoxy putty then further
secured with a 2 inch band of carbon fiber running the length of the root chord. This will not
only mitigate the point of failure that is where the fins attach, but it will also reduce the
coefficient of drag on the rocket.
The point of failure is the airframe. The airframe will be prone to bending and flexing
during the rockets flight. However given that the body tube is composed of fiberglass it is not
likely to bend or flex. This is merely an assumption so in order to properly assess the risk of
failure when it comes to the body tube we did research followed by calculations. Our research
was that of looking into the Contest - Roc Yahoo group and using their Flex- Roc calculator in
order to plug in our rockets components and materials to predict whether or not it will flex and
break during its flight. The program ran many iterations of the formulas involved to predict the
materials failure during flight every step of the way. It was solving for an aeroelastic divergence.
The software found that our rocket was not in fear of aeroelastic divergence as the team had
predicted, and therefore the rocket was not at risk of flexing to a point of destruction. So this
point of failure was properly mitigated by picking a strong enough material to house the rockets
internals for flight.
The next point of failure is the recovery system. will be as follows. To start with the
chute material will be a Rip Stop nylon with a rating of 70DEN FR to be provided by Seattle
Fabrics. The FR designation means fire retardant, as the chute is a polyurethane coated fabric.
With the chute diameter measured out to eighty-four inches we will be using an eighth inch Flat
Braid Dacron which is rated up to four hundred pounds. It can be found at paragear under the
category number W9754FR. We chose the eighth inch Dacron over the three sixteenth inch as it
is lighter and conserves space this way. Also the eighth inch Dacron is strong enough as each
line is rated up to four hundred pounds. So with there being eight gores this gives our shrouds the
cumulative strength of thirty-two hundred pounds, or about 320G's. Meaning it would take an
opening shock at 320G's of force to cause a failure in the chute material or shroud lines, which is
extremely unlikely to happen at it's time of deployment. Also as far as the shroud line goes we
are using a flat fabric as opposed to round due to it being easier to sew with our given skill set.
Two types of thread line will be provided by Seattle Fabrics for use in the chute. The first being
Z46 thread used to sew the gores together. It is size forty-six and one ounce thread that is nylon
bonded giving it the Z designation. The next being the Z69 to sew the gores and threads to the
shroud lines. It is size 69 and nylon bonded giving it the Z designation. As mentioned prior we
plan on using eight gores in the construction of the chute. The gore length is fifty inches for this
semi hemispherical design with a gore width of thirty three inches. It has an apex of forty-five
and base angle of sixty. The base angle is derived from the fact that a semi hemisphere is two
thirds of a full hemisphere leaving us with an angle of sixty instead of ninety. Using loops we
will be able to give each piece two shroud lines sewn the length of the gore giving it reinforced
strength. This also makes it easier to hook the parachute to the hardware as it will be presented in
a loop instead of leaving us with shrouds to knot. We will be running French fell seams along the
entirety of the chute which is a strong stitching pattern. This clearly has been determined to be a
strong pattern for our main chute being the semi hemispherical one. The X-form drogue has
already been sewn in house with all materials and techniques aforementioned. It has so far
successfully brought down a level two rocket flying a J class motor that weighed in at ten pounds
with a height of eighty four inches.
Next we have the hardware. The shroud lines being composed of loops will be attached to
to a D Link. This D link will come from a standard hardware store and is size five sixteenths.
This gives it a strength rating of 1500 pounds or 150G's. The D Link will hook into a barrel
swivel that is provided by LOC precision. It is swivel 12 and rated for 1500 pounds of strength.
Which is again at 150G's of force in order to induce failure. The barrel swivel will be hooked
onto our quarter inch tubular kevlar shock cord which carries a 3000 pound strength rating.
Which comes down to 320G's of force to induce failure. The shock cord will be tied onto a Ubolt that will be anchored in a carbon fiber bulk head made with 24KPL50 which provided by
solarcomposites. According to their website it boasts a rating of aerospace 600KPSI.
Through the parachute design and hardware applications the point of failure is
determined to be the D-link or barrel swivel which are rated for 150G's of force or 1500 pounds.
However seen as at apogee the rocket will not nearly be under this pressure and according to the
rate of descent calculator once the drogue is deployed it will descend at 24 feet per second.
Meaning even when the main chute deploys it will still not experience 150G's of force, thus
giving our chutes a theoretical point of failure at the D-Link to barrel swivel connection however
it is not likely to happen. It has been properly mitigated by analyzing all possible points of failure
then looking at the feasibility of it actually happening.
The next point of failure during the failure analysis is the electronics components. The
electronics have three points of failure. First they can fail due to improper wiring, which would
cause the electronics to malfunction and not work at all. Second they can fail due to an improper
power source, without power they cannot run. Finally they can fail due to manufacturer defect. In
order to mitigate against these points of failure in the electronics, they must be thoroughly
inspected for the faults. Then the appropriate action must follow, whether it is changing batteries
or adjusting wiring. It can be difficult to mitigate for a manufacturer defect however testing upon
receiving the components should work well towards combating it.
Another common point of failure is the nose cone. As the nose cone is held onto the
rocket by a shock cord, special care must be taken to mitigate the possible of a failure of the nose
cone to be held to the rocket. The team is attempting to mitigate this risk by using a nose cone
formed out of a kevlar and fiber glass composite to give it structural rigidity, as well as attaching
it firmly to the shock cord using strong materials such as metal bolts and metal wire. This will do
well to prevent a point of failure in the nose cones attachment to the rocket.
During the failure analysis the boat tail has also come under scrutiny. The boat tail could
present a potential point of failure if it is not firmly attached to the rocket. In order to mitigate
this the boat tail will be attached to the rocket by using an epoxy resin. Epoxy resin should be
strong enough to hold the boat tail to the rocket as the boat tail isnt taking a large force of
pressure during flight.
The rockets over all safety has been determined by this failure analysis. The safety of the
rocket will be quite assured once the rocket is assembled and ready for flight. Additional safety
hazards the rocket may be concerned with outside of the aforementioned failure points, are it’s
shipping and transportation to the launch site. Transportation is a big logistic of such a large
rocket and as such it is mission critical that no parts be damaged during its transport. This
presents a large concern for the rockets safety however there is not much the team can do to
mitigate this safety concern if the team decides to have the rocket shipped as it is entirely up to
the people who handle the rocket during its transit. Although the team can take preventative steps
to package it to try and reduce and damage due to mishandling during transport.
How the subscale flight data has impacted the design of the full-scale launch vehicle.
Results from the subscale launch will determine if the set up of the electronics such as the
accelerometer to the arduino, the strain gauges to the arduino and Raspberry Pi, and cameras to
the Raspberry Pi are set up correctly to record the appropriate data. This includes appropriate
coding of the Arduino, and Raspberry Pi. Data will be analyzed from the subscale launch to see
if the data makes sense and adjustments will be made accordingly.
Recovery Subsystem
Beginning with the parachute the specifications will be as follows. To start the chute
material will be a Rip Stop nylon with a rating of 70DEN FR to be provided by Seattle Fabrics.
The FR designation means fire retardant, as the chute is a polyurethane coated fabric. With the
chute diameter measured out to eighty-four inches we will be using an eighth inch Flat Braid
Dacron which is rated up to four hundred pounds. It can be found at paragear under the category
number W9754FR. We chose the eighth inch Dacron over the three sixteenth inch as it is lighter
and conserves space this way. Also the eighth inch Dacron is strong enough as each line is rated
up to four hundred pounds. So with there being eight gores this gives our shrouds the cumulative
strength of thirty-two hundred pounds, or about 320G's. Meaning it would take an opening shock
at 320G's of force to cause a failure in the chute material or shroud lines, which is extremely
unlikely to happen at it's time of deployment. Also as far as the shroud line goes we are using a
flat fabric as opposed to round due to it being easier to sew with our given skill set. Two types of
thread line will be provided by Seattle Fabrics for use in the chute. The first being Z46 thread
used to sew the gores together. It is size forty-six and one ounce thread that is nylon bonded
giving it the Z designation. The next being the Z69 to sew the gores and threads to the shroud
lines. It is size 69 and nylon bonded giving it the Z designation. As mentioned prior we plan on
using eight gores in the construction of the chute. The gore length is fifty inches for this semi
hemispherical design with a gore width of thirty three inches. It has an apex of forty-five and
base angle of sixty. The base angle is derived from the fact that a semi hemisphere is two thirds
of a full hemisphere leaving us with an angle of sixty instead of ninety. Using loops we will be
able to give each piece two shroud lines sewn the length of the gore giving it reinforced strength.
This also makes it easier to hook the parachute to the hardware as it will be presented in a loop
instead of leaving us with shrouds to knot. We will be running French fell seams along the
entirety of the chute which is a strong stitching pattern. This clearly has been determined to be a
strong pattern for our main chute being the semi hemispherical one. The X-form drogue has
already been sewn in house with all materials and techniques aforementioned. It has so far
successfully brought down a level two rocket flying a J class motor that weighed in at ten pounds
with a height of eighty four inches.
Next we have the hardware. The shroud lines being composed of loops will be attached to
to a D Link. This D link will come from a standard hardware store and is size five sixteenths.
This gives it a strength rating of 1500 pounds or 150G's. The D Link will hook into a barrel
swivel that is provided by LOC precision. It is swivel 12 and rated for 1500 pounds of strength.
Which is again at 150G's of force in order to induce failure. The barrel swivel will be hooked
onto our quarter inch tubular kevlar shock cord which carries a 3000 pound strength rating.
Which comes down to 320G's of force to induce failure. The shock cord will be tied onto a Ubolt that will be anchored in a carbon fiber bulk head made with 24KPL50 which provided by
solarcomposites. According to their website it boasts a rating of aerospace 600KPSI.
Through the parachute design and hardware applications the point of failure is
determined to be the D-link or barrel swivel which are rated for 150G's of force or 1500 pounds.
However seen as at apogee the rocket will not nearly be under this pressure and according to the
rate of descent calculator once the drogue is deployed it will descend at 24 feet per second.
Meaning even when the main chute deploys it will still not experience 150G's of force, thus
giving our chutes a theoretical point of failure at the D-Link to barrel swivel connection however
it is not likely to happen.
Electrical components and how they will work together to safely recover the launch vehicle.
The main electrical component in charge of the recovery of the rocket is our PerfectFlite
stratologger. The PerfectFlite stratologger will be wired up to a terminal block which will then be
wired to two ejection charge cups made of PVC. Through previous experience the PVC cup and
terminal block will work to deploy the ejection charge contained within the cup after carrying the
charge from the flight computer. The flight computer and terminal block will be wired with size
eighteen AWG running between them all. It can handle up to three hundred volts which gives it
full capability to carry the charge from the PerfectFlite stratologger to the terminal block. So it is
a simple system that has the flight computer working with a terminal block and charge cup all
brought together through wiring to cause black power charges to go off and deploy the recovery.
The electronics system for the deployment system is a simple as a flight computer sending a
charge through a circuit. Possible points of failure through this design lay within the wiring or
battery. In order to mitigate these points of failure to next to nothing it is simple as making sure
all wiring is secure through the ports connected, and that the battery is holding the proper amount
of charge. Assuming this the flight computer will easily be able to send its charge to the ejection
cups without failure.
Include drawings/sketches, block diagrams, and electrical schematics.
Raspberry Pi System Schematic
Arduino Uno System Schematic
Kinetic energy at significant phases of the mission
The major changes in the kinetic energy of the mission occur at launch, completion of
motor burn, apogee, drogue deployment, main deployment, and landing.
Test results.
As previously discussed the parts for the subscale and full scale have not yet arrived so
testing has been limited to ground testing utilizing equipment and parts that were already
available to the team. Results of the ejection charge tests showed that two, 1.5 g black powder
ejection charges per break would be enough to ensure the safe deployment of the recovery
system if one of the charges failed due to faulty wiring, a faulty altimeter, or incorrect arming of
the black powder charge.
Safety and failure analysis.
The safety and failure analysis of the recovery system hinge on the components it is made
out of. To begin with the barrel swivel and D-link have a strength rating of 1500 pounds or
150G’s. This makes them a failure point as if there is more than 150G’s of force exerted on the
rocket this could cause a hardware failure. However the rocket is unlikely to encounter more than
150G’s of force during its flight. The next point of failure during analysis of the recovery system
is the shock cord. The shock cord could potentially burn or snap however as it is a ¼ inch tubular
kevlar shock cord, it is flame resistant to 2000 degrees fahrenheit and is also capable of
withstand up to 3000 pounds of tensile strength, or 300G’s. As it is highly unlikely to experience
either of those conditions during flight it should be safe. The next point of failure may be the
parachute. With the various cords and threads being described in detail above it is easier to just
note the point of failure for the parachutes will be approximately 3200 pounds of tensile strength
due to its quality materials. This actually makes the parachute the strongest point of this failure
analysis as it has the capability to withstand 320G’s. There is also the matter of the anchor of
what the shock cord is anchored to. The shock chord is anchored to a U bolt that is mounted
inside of a carbon fiber bulkhead. This creates an additonal safety hazard during this failure
analysis. However it is being mitigated by being made of carbon fiber which according to the
manufacturer solarcomposites can withstand up to 600000 kpsi of force. Which should prove
more than adequate for withstanding any force exerted upon it by the shock cord. Through this
the hardware pieces of the recovery system have been noted for their various points of failure.
However due to the strength of all these materials it is easy to say the potential for failure has
been mitigated drastically by picking materials that are able to withstand these forces acting upon
them.
The PerfectFlite Stratologger is what causes the recovery system to deploy so it is also a
crucial point of failure. Should the flight computer fail to deploy the recovery system, all of the
strength of the materials ceases to matter as they are unable to do their job. In this case the proper
mitigation for this scenario would be to test the flight computers before being installed, as well
as once installed. It would also be keen to check the points of failure of the flight computer
which includes the wiring and power source. Once the power source has been verified as charged
and the wiring declared functional, the flight computer should work. Mitigation for those two sub
components is as simple as testing the current to ensure that enough power is being supplied.
Once this is done and the flight computer passes pre flight checks, it is safe to deploy the
recovery system.
Mission Performance Predictions
Mission performance criteria.
A successful mission involves all systems and subsystems working properly. The rocket
itself must reach the desired altitude of 8,500 feet and the recovery system must deploy to let the
rocket descend safely to the ground. The first of the payload criteria, the hazard detection system,
involves the camera capturing a live video feed which must be analyzed by the software on the
Raspberry Pi microcontroller. The feed and hazard detection overlay must be transmitted to a
computer at the ground station in real time. The strain gauges, accelerometers, altimeters, and
voltmeter, all must function completely in order to collect data that will meet the criteria for the
analysis in payload option 3.2.1.3. Aerodynamically, the rocket must perform as closely as
possible to the specifications and coefficients of drag determined using computational fluid
dynamic simulation and wind tunnel analysis.
Flight profile simulations, altitude predictions with final vehicle design, weights, and actual
motor thrust curve.
Thrust Curve courtesy of ThurstCuver.Org
Show stability margin and the actual CP and CG relationship and locations.
Payload Integration
The launch vehicle will have two electronics bays, one in the hollow nose cone, the other in
the mid section of the rocket. The reasoning of using two electronics bays is due to the placement
of two separate strain gauges. The team’s objective is to measure the strain endured by the
launch vehicle during launch in two separate locations, therefore, two electronics bays are
needed. The electronics bays will be enclosed cylinders and double as couplers for the
rocket.The electronics bay in the nose cone will contain an accelerometer, a sonic beacon, a
battery, and a Raspberry Pi which will be connected to a camera located just outside the nose
electronics bay at the base of the nose cone. The camera will be spring loaded against the inner
side wall of the structure of the vehicle. The camera will be designed to slide out and deploy as
the nose cone is ejected from the rest of the rocket during the deployment of the main parachute.
The Raspberry Pi will also be wired to a strain gauge on the outer part of the base of the nose
cone by drilling small holes through the lower part of the nose cone to feed the wires into the
electronics bay.
The second electronics bay will be in the mid section of the rocket roughly 39 in. from the tip
of the nose cone. This electronics bay will house the altimeter, Arduino, accelerometer, and two
9V battery. The outer ends of the electronics bay will have a small PVC cup on either side that
will hold black powder and will serve as the ejection charges for the recovery system. The
Arduino will also be wired to a strain gauge on the outside of the structure of the rocket in the
same way as the nose cone electronics bay.
Ease of integration
Integration plan.
Payload will be integrated and stabilized using a center panel bolted in the electronics bays.
The Arduino components, transmitter, altimeter, accelerometer, Raspberry Pi, batteries, and
cameras will be stored on one either side of the center panel. This will be done to ensure that
none of the contents of the rocket shift haphazardly during launch. The electronics bays will be
enclosed cylinders that also serve as couplers for the rocket and will be easily compatible with
the rocket. The electronics bay bulkheads will be able to be screwed on and off and the
electronics will be on a sled capable of easily being removed and inserted into and out of the
electronics bays. Ejection charges will be mounted on the ends of the electronics bays which
eliminates the probability of them moving during flight to ensure a safe deployment of the
recovery system. The most difficult assembly of this launch vehicle will be leading wires
through the structure of the rocket to the strain gauges. Only a very small portion of wire will be
exposed to the outside of the launch vehicle to reduce the possibility of the strain gauges
shearing off during flight. What little wire that will be on the outside will be taped down in such
a way as to make the strain gauge wire as streamlined as possible.
Installation and removal, interface dimensions, and precision fit.
Installation of the parts will be simple. First the electronics bay bulkheads will be undone.
This will allow us to pull the length of wire from the switches for the altimeter and black powder
charges, and wire them onto the perfect flite. Then we can wire the 9 volt to the perfectflite. This
is of course taking into account that the perfectflite is mounted to an internal sled that is
composed of a wooden slab with rail guides on the back of it. The additional electronics for the
ebay will also already be pre mounted to the wooden slab. They will also be wired to the
internals they respectively control, such as the strain gauges. Once this is done the sled will be
put on the internal steel rod of the electronics bay and slid into position amidst all the wires. The
bulkhead will be resealed and thus the installation of the mid electronics section is complete.
For the installation of the nosecone the process is largely the same when it comes to the slab
being inserted. However the slab is not to be in place by a rod inside the nose cone. Instead it is
installed by simply inserting the sled to fit into centering rings inside the nose cone. Then a
bulkhead will be placed at the bottom of the nose cone with a bolt placed horizontally beneath it
to hold it in place. This bolt will have a length of steel wire then tied onto it then the kevlar shock
cord respectively. Thus the nose cone electronics bay is installed.
Removal of electronics is as simple as repeating the process in reverse. Bulkheads are
undone, then the internal wires are disconnected and the sled is slid off of the rod it was resting
on in the internal space.
Interface and precision fits for the electronics have been taken into consideration. It will be a
precision fit as all pieces will fit in an exact location with consideration to its job and space
available. They will interface with other pieces according to what they have designated to do
when it is time to perform. E.g the flight computer with the ejection charges and the
microprocessors with the sensor payload.
Compatibility of elements
All of the elements of the payload will work together to accomplish the objectives of the
mission. The 5MP camera module, Raspberry Pi, and 3G Dongle receiver will work together to
give the team the ability to monitor the rocket in flight. The strain gauges, voltmeters, and
accelerometer will work together to allow the team to monitor the strain, electrical system, and
propulsion of the rocket in flight. The elements will all be placed inside of the body tube and
they will be kept apart to prevent damage and experimental errors.
Simplicityof integration procedure.
Components were selected for their simplicity in integration into the electronics bay. The
Raspberry Pi system was specifically designed to operate easily. The Raspberry Pi’s 5MP
camera was chosen for its easy integration with the Raspberry Pi as it was manufactured
specifically for its use. The pi was also picked for the simplicity of integrating the 3g dongle to it
as well. Since so many repositories of code and tools exist for the raspberry pi it is possible to
easily integrate all of its peripherals into its operations.
Launch concerns and operation procedures
Recovery preparation.
Recovery preparation will begin with taking the electronics bay out from the mid section
of the rocket, and then loading the appropriately sized black powder charges into their cups on
their respective sides. Then the black powder charges will be secured into the cup and the mid
section will be reloaded. Once it is reloaded the parachutes will be hooked into their anchors
inside the rocket and then packed in a manner conducive to an easy release upon ignition of the
ejection charges. Once this is complete the rocket will be reassembled and any shear pins or
screws will be installed in their respective locations. After this the altimeters pull pin will be
removed turning it on. After the readout of data is output by the altimeter, the switches for the
black powder charges will be turned to the on position, allowing them to be deployed when the
charge comes from the altimeter. After this is all done the system will be in an armed state and
the recovery system will be ready for its deployment.
Motor preparation.
The motor will be prepared by first removing the K660 out of its shipping package. The
motor will then receive an inspection for any flaws that could be present due to manufacturer
error. Once the motor has passed this inspection it will then be covered in lithium grease so that
the motor may be then slid into the motor casing. The lithium grease also prepares the motor not
only for launch, but for post recovery cleaning as it allows the motor to be removed from the
casing easily. Once the motor is slid into the casing, the bottom ring of the casing will be
secured. Then the casing will then be slid into the motor mount and then secured by the Aeropak
motor retainer. Once the retainer ring is secured tightly onto the retainer the motor will be
deemed secure in the rocket. Then following this a pyrogen based electronic igniter will be
inserted into the motor and then secured onto the power source with special consideration taken
to not touch the leads together. At this point everyone will retreat away from the pad to the
minimum safe distance in order to observe the launch. All work will be considered acceptable if
the range safety officer approves, as he has the finals say.
Igniter installation.
As mentioned above the igniter will be installed by being inserted into the fully prepared
motor once it is on the launch pad. It then will be hooked up to the power source with special
consideration taken to prevent the leads from touching. Once this is done the igniter will be
deemed ready.
Setup on launcher.
In order to be set up on the launch rail, the rocket must have the motor and recovery
system properly installed as outlined above. Then the rocket will be slid onto the launch rail by
its rail mounts. Then the igniter will be installed as it is stated above. Then rocket will be then
deemed to be setup on the launcher.
Troubleshooting.
Trouble shooting at launch will be mainly concerned with the failure of the igniter or the
onboard launch components of the rocket. If an onboard electronics component is failing the first
troubleshooting step will be to check the connections in order to make sure power is able to flow.
If the connections are valid, then the power source will be checked for power. If that too has
failed then troubleshooting will involve a more direct examination of the electronics to see if
there is any damage to the piece causing it to malfunction. Troubleshooting with the igniter will
be to see if it is improperly hooked up to the leads causing it to not ignite the rocket, or that the
power source needs to be examined in the event the rocket doesn’t launch upon the pressing of
the ignition button. These are the initial steps of troubleshooting two of the most probable points
of failure but first beginning with a simple examination.
Postflight inspection.
Afterward the rocket is recovered, the team will inspect the launch vehicle for any critical
damage. This will include but not be limited to the body tube, fins, internal electronic
components and wiring, power systems, and measurement devices, among others.
Safety and Environment (Vehicle)
PreEffect(s)
RAC Mitigation
electrical burns,
skin damage,
Careful grounding of
damage to
electronics and
Testing flight
electrical
persons working on
computer and
components, start
electronics, safety
Exposure sensors; Installing a fire to
personell in vicinity,
to electric flight computer and surrounding work
use certified
shock
sensors
area
3D components
fire suppressant easily
Burns, skin
accessible,
injuries, lung
masks/open air when
irritation,
working with powder,
damages to
safety officer present,
airframe and
maintain minimum
testing/installing
electronics bay,
safe distance of 10 ft
Exposure recovery charges cause a fire to
when testing and have
to black
and motor ignition surrounding work
no personell in line of
powder
system
area
3C sight of firing direction
wear masks and
gloves when working
with epoxy and epoxy
related materials, have
safety officer present,
construction of the skin damage, lung
work in an area with
vehicle structure, damage, damage
adequete ventilation,
Exposure payload mounts,
to airframe, highly
avoid heat and spark
to epoxy
and motor mounts flammable
3C sources
constrution and
skin irritation, lung
wear masks and
Exposure machined
irritation, damage
gloves when working
to
alterations to
to electrical
with fiberglass and
fiberglass airframe and
components if
fiberglass related
dust
related structure
exposed
2D materials
construction and
having trained team
machined
members use
Machining alterations to
machining equipment,
related
airframe and
cuts, skin injuries,
wearing safety glasses
hazards
related structure
eye injuries
D2 and general PPE
Hazard
Cause
Identify safety officer for your team.
Verification
PostRAC
Airframe
Loading
Procedure 2.1.1
D44D
- Safety and
Handling of
Electronics Bay
Components
Airframe
Loading
3D
Procedure 1.1.2
- Safety and
Handling of
Recovery
Charges
UCF Machine
Lab Composite
Materials
Procedure
UCF Machine
Lab Composite
Materials
Procedure
UCF Machine
Lab-Machine
Lab Safety
Procedure
3D
3D
3D
The safety officer of the team is Anthony Laiuppa. He has been picked based on his
qualification of holding a level two high power rocketry license through the National Association
of Rocketry for the past two years. This along with his experience with high power rocketry and
the regulations around them gives him a strong base of knowledge involving not only the hazards
of construction the rocket, but also the regulations surrounding high power rocketry. He has
previously led and supervised two previous SEDS sound rockets produced by the club to to
compete in the SEDS high power rocketry competition 10,000 foot challenge. He will be
responsible for the safety and regulation of the team during construction and launch.
Risk
Consequence
Mitigation
Systems or components fail
during testing and/or test
launches.
Systems and components
may have to be
redesigned or replaced.
Order extras of any possibly fragile
components to prevent
manufacturing delays. Have backup
ideas for any systems that may fail.
Parts arrive late.
Manufacturing, testing,
and test launches are
delayed.
Order parts well in advance and
ensure that they are in stock.
The project costs more than
the team estimates.
Project is over budget
and team may not have
sufficient funds to
continue.
Constantly maintain and update
budget and overestimate costs if
possible.
The project takes longer to
complete that the team
estimates.
The team is behind
schedule.
Keep all team members informed of
the planned schedule and start
working on aspects of the project
earlier than necessary whenever
possible.
The rocket does not meet
RSO requirements on
launch day.
The team loses points
and does not get to fly.
Work closely with NAR Team
Mentor to verify that the rocket is
being built to the required standards.
The team loses members
Project is severely
and does not have sufficient behind schedule and may
personnel to complete all
not be completed.
aspects of the rocket.
Prioritize work in order to meet the
most significant criteria.
List of personnel hazards and demonstration research is as follows.
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ROL ATF News
Federal Explosives Law (790k)
Exemption Paper Trail (613k)
List of Explosive Materials (96k)
Federal Firearms Laws (227 k)
Letter from ATF 1-13-99 (57k)
AeroTech to J. Zamillo 11-4-98 (70k)
ATF Communique 4-25-94 (53k)
Memorandum of Opinion 6-24-02 (1.48MB)
Letter From ATF 4-20-94 (80k)
Letter From ATF 6-20-94 (66k)
G. Brunda to ATF 5-28-92 (104k)
ATF to G. Brunda 6-5-92 (21k)
ATF to T. Smith 11-28-01 (35k)
GCR ATF Affidavit 6-27-02 (217k)
Court Order re: TRA/NAR vs. ATFE 3-16-09
Discuss any environmental concerns.
The environmental concerns that the team anticipates are largely confined to the specifics
of the Bonneville Salt Flats. To begin with possible environmental concerns start with the
weather and conditions of the launch site. Rain is an environmental concern that carries the
capability to possibly damage electronics. There is also that of salt of dust which are concerns for
an area as arid as the salt flats as they can potentially damage electronics through corrosion or
interference. The rocket also faces the environmental concern of the ground. The salt flats have a
rough compact surface which can be potentially damaging to the rocket depending on its rate of
descent as it lands on the surface. Also the rocket carries the potential of landing near hazardous
indigenous wildlife that is native to the environment present. In the event of this, it could also
pose a threat to the team during the retrieval process of the rocket. Also the team faces the
environmental concerns faced with human health in a dry environment. This means proper
hydration as well as taking precautions to make sure no one injures themselves through accidents
on the flats. Such as falling and scraping their knee as an obvious example. That is a viable
environmental concern as rough ground is extremely hazardous towards fragile human skin.
Simple mitigations towards these environmental concerns are as easy as taking the obvious
proper precautions such as bringing water or avoiding the native wildlife.
IV) Payload Criteria
Testing and Design of Payload Experiment
Design at a system level.
Drawings and specifications
Analysis results
Analysis of the payload involved making sure each part used in the experiment worked
together, analyzing which components would be best suited for our experiment and finding
which methods would be most cost-effective in creating our experiment. We found that the
Raspberry Pi system would fit all of the requirements for our payload experiment.
Test results
Tests resulted in a successful performance of our payload components. Cameras,
transmitters and data collection performed successfully according to our payload missions.
Integrity of design
The design of the system has been carefully implemented to ensure the integrity of all
components and systems. The design has been created in order to collect and transmit accurate
and precise data.
#
Requirement
Feature
Verification
3.1.1
The payload shall
incorporate a camera
system that scans the
surface during descent in
order to detect potential
landing hazards.
The rocket payload will include a
camera that provides a video feed
during descent.
The camera will be
tested during half
scale and full scale
launches.
3.1.2
The data from the hazard
detection camera shall be
analyzed in real time by a
custom designed on-board
software packaged that
shall determine if landing
hazards are present.
The rocket will contain a
Raspberry Pi microcontroller with
original software that will analyze
the video feed for hazards.
The software will
be tested for hazard
locating accuracy
during the half
scale and full scale
launches.
3.1.3
The data from the surface
hazard detection camera
and software system shall
be transmitted in real time
to a ground station.
The payload will feature a 5.8 Ghz
lightweight video transmitter to
send the feed and hazard analysis
to a computer on the ground.
The transmitter will
be tested for range
and quality during
the half scale and
full scale launches.
3.2.1.3 Structural and dynamic
analysis of air-frame,
propulsion, and electrical
systems during boost.
Strain gauges will be placed on the
body tube to measure strain. The
electrical system will be analyzed
by voltmeters. An accelerometer
and altimeter will be used to
measure the acceleration and
velocity of propulsion. The
Arduino Uno will gather data.
These systems will
be tested during the
half scale and full
scale launches.
3.2.2.2 Aerodynamic analysis of
structural protuberances.
The team will use computational
fluid dynamics simulation in
SolidWorks and wind tunnel
testing to analyze the
aerodynamics of the rocket.
The data gathered
from the analysis
will be verified
during the half
scale and full scale
launches.
Workmanship as it relates to mission success.
The teams approach to workmanship will be following part of the UCF Creed of
Integrity, Community, Creativity, Scholarship and Excellence. In this case Team Advance will
be specifically focusing on the word Excellence. As a team we would like to excel in
workmanship as we represent UCF through our attempt at mission success. It is not only the best
way to represent UCF by following the UCF creed, but also the best way to chase after mission
success through workmanship by having excellence in all the team does. This means working
hard to insure that the rocket isn't just adequate to achieve the mission, but that it excels in all it
does. It is vital that all tasks are completed using a high level of precision as to avoid making
mistakes that would compromise the functionality or the scientific objectives of the mission.
Planned component testing, functional testing, or static testing.
The team has planned component testing that will be performed once the ordered parts
arrive. The team will be testing all of the electronics of the vehicle prior to launch as well as
during the half scale and full scale launches. The team will be testing the functionality of the
rocket using Rocksim as well as testing the actual rocket during the launches.
Status and plans of remaining manufacturing and assembly.
Project Advance is currently awaiting the arrival of parts and will begin manufacturing
and assembly once they are received.
Describe integration plan.
Integration of payload has been planned to assure that instruments collect accurate data.
The Arduino Uno as well as the Raspberry Pi will be used to capture data taken by sensors. The
Arduino Uno will act as our flight computer taking input from strain gauges. Data will be taken
continuously for set intervals during flight and stored on the Arduino Uno. The Raspberry Pi will
be integrated with camera, 3G Dongle, accelerometer and strain gauges.
Precision of instrumentation and repeatability of measurement.
All components of the payload experiment will be programmed and calibrated to ensure
maximum precision during measurement. The software systems on board both the Raspberry Pi
and Arduino Uno will be capable of gathering data during multiple tests. Data will be transferred
to computers in order to free up memory on these microcontrollers so that they can gather more
data.
Payload electronics with special attention given to transmitters.
Nose Cone Payload:
Nose cone payload consists of the Raspberry Pi system with camera, batteries, and 3G
Dongle, a breadboard and an accelerometer. The Raspberry Pi will transmit live video feed with
landing hazards to the ground station. Connected to the Raspberry Pi will also be a breadboard
with strain gauges and an accelerometer to acquire data during flight.
Electronics Bay Payload:
The electronics bay payload contained in the body tube will consist of a Arduino Uno R3
Microcontroller.A strain gauge will also be attached to the Arduino Uno to collect data on the
strain put on the rocket’s frame.
Drawings and schematics
Electronics Bay Payload Schematic
Nose Cone Payload Schematic
Electronics Bay Payload Block Diagram
Nose Cone Payload Block Diagram
Batteries/power
Item
Battery Type
Raspberry Pi Model B
10,000mAh lithium ion battery
Arduino Uno R3 Microcontroller 9V battery
Transmitter frequencies, wattage, and location
Transmitter
Frequency Wattage Location
Raspberry Pi 3G Dongle 2.4 Ghz
5 Watts
Nose Cone
Test plans
Tests for transmitters include the following
Connecting the 3G Dongle to the Raspberry Pi and ensuring that Internet connection can be
achieved. Scripts will be run on the Raspberry Pi to confirm if an Internet connection can be
achieved.
Safety and failure analysis.
Failure in the transmission components of the payload will compromise the success of the
payload mission. Failure of nose cone payload may happen in the camera component, the 3G
component, strain gauges, accelerometer, breadboard, and the Raspberry Pi itself.
The camera’s failure would result in not being able to receive live video stream of the ground
hazards during descent. Failures would be caused from the camera failing to deploy, the camera
becoming damaged due to corrosion from the salt in the environment, or the camera becoming
disconnected from the Raspberry Pi. Failure of the 3G dongle from the Raspberry Pi would also
prevent hazards from being able to be detected. The dongle may become detached from the
Raspberry Pi during flight or fail to connect to the Internet.
The failure of the breadboard would result in a loss of data from components acquiring
information during flight. Failure in the breadboard could be caused by faulty wiring or the
breadboard becoming disconnected from the Raspberry Pi. Without these components
functioning properly, the system would not be able to collect flight data, resulting in a failure of
the payload objective.
Payload Concept Features and Definition
The purpose of the requirement 3.1 is to develop a camera system and software that will
scan for landing hazards that will support the development of future SLS technology. This
simplified system will explore a potential hazard identification method that could be used on a
full-scale mission. The camera and software systems will be the original work of the team.
The exterior of the rocket will be covered in mylar, thus accentuating the body of the
rocket and increasing visibility. The nose cone is a Von Karman which will produce minimal
drag on the rocket and the fins and nose cone will be composed of 100 percent fiberglass. The
drogue chute will be x-shaped which will be able to slow the rocket more efficiently than
streamers and will be more effective than a parachute.
One of the greatest challenges the team faces is the lack of experience in the area of
rocketry among its members. The group is composed of underclassman, including two freshmen,
who have not partaken in any rocket launches or engineering projects. Options 3.2.1.3 and
3.2.2.2 involve advanced stuctural, dynamic, and aerodynamic analysis of the entire launch
vehicle. Combined with 3.1 and the lack of experience among members, these objectives will
prove to be a suitable challenge for undergraduate students which meets the intention of this
competition.
Science Value
Payload objectives.
The payload objectives include monitoring flight data, including acceleration, altitude, and
strain on the rocket frame. Data will be collected continuously during flight at a set interval.
Also, live video transmission must be received by the ground station during descent.
Payload success criteria.
Payload success is defined by the following criteria:
o Payload must be able to be recovered and reusable
o Flight computer sensors collect accurate data during flight and is collected after
landing
o Hazard detection sensors must transmit live, accurate data on landing hazards
during landing
Experimental logic, approach, and method of investigation.
The logic behind this experiment is to collect data that will be helpful in the creation of
future full-sized rockets. The experiment is being approached from an engineering perspective,
with the rocket assembly being the most important component of the project. Ensuring that all
parts of the rocket will be operational at all stages of the launch is vital to the collection of data,
and thus extremely important to the overall success of the mission. The data collected from the
launch will be significant for future rocket launches, making this a scientific experiment as well.
The experiment is being investigated through careful planning and controlled testing through
Rocksim, SolidWorks, and practice launches.
Test and measurement, variables, and controls.
The testing of the vehicle will be conducted through the computer softwares Rocksim and
SolidWorks and the test launch that the team plans on having closer to the launch date. Through
test launches the team will measure acceleration, velocity, and strain. The team will be
calculating thrust and impulse, measure the altitude the rocket travels, analyze the recovery
system, and the aerodynamics of the rocket.
Relevance of expected data and accuracy/error analysis.
The expected data is meant to help develop future rockets and spacecraft. The data the team
is expecting will confirm that all of the payload criteria are met and it will allow future Student
Launch teams to improve upon ideas and procedures of this project. Surveying the ground for
potential landing hazards will help research methods of avoiding hazards in real world
applications such as for future spacecraft or even drones. Measuring strain on the structure of the
launch vehicle will assist in the building of future launch vehicles by demonstrating whether or
not the fiberglass used for this rocket is able withstand the forces of launch. Measuring
acceleration and velocity to calculate thrust and impulse will either verify or disprove the
manufactures specifications of the motors the company produces. It will also demonstrate a
means of calculating thrust and impulse as a backup method for other vehicles if the primary
source of measuring thrust and impulse is lost. Possible errors include faulty wiring, incorrect
coding of the Raspberry Pi or Arduino, improper loading of the motor, improper folding of the
parachutes, and uncertainty of the Raspberry Pi, Arduino, strain gauges, accelerometer, and
altimeter.
Experiment process procedures.
Prior to launch the experimental payload will be prepared and checked to ensure the
system is operational. Once it is confirmed the system is loaded the team will take readings on all
of the payload systems in order to establish a base reading. During the flight the experiment will
proceed as the payload continues to take readings with its sensor as the rocket climbs in altitude.
As the rocket reaches apogee and begins to descend the cameras will be deployed continuing on
the experiment as they begin to stream video in real time so that hazards may be located. The
data will then be analyzed in real time. Once the flight is over the team will recover the rocket
and remove all experiment equipment. It will then proceed to power the equipment on and
retrieve the data that has been recorded to it. Once the data is collected, the team will complete
analysis to determine if the payload criteria were met and outside factors will be considered in
the conclusions reached by the group.
Safety and Environment (Payload)
The safety officer of the team is Anthony Laiuppa. He has been picked based on his
qualification of holding a level two high power rocketry license through the National Association
of Rocketry for the past two years. This along with his experience with high power rocketry and
the regulations around them gives him a strong base of knowledge involving not only the hazards
of construction the rocket, but also the regulations surrounding high power rocketry. He has
previously led and supervised two previous SEDS sound rockets produced by the club to to
compete in the SEDS high power rocketry competition 10,000 foot challenge. He will be
responsible for the safety and regulation of the team during construction and launch.
Failure analysis: The failure analysis begins with the outermost layers and then proceeds
inwards. The failure analysis for the airframe can be broken down into three parts. The point of
failure regarding the stability of the rocket, the point of failure regarding its fins, and the point of
failure that is the body tube itself.
The failure analysis that has been done on the stability of the rocket has been modeled in
RockSim and the data has been collected accordingly. Through RockSim we calculated the
center of pressure and center of gravity using the advanced Barrowman equation that was
developed by NASA engineer James Barrowman. Standard static stability margin of 2 body
diameters is the industry accepted safety margin used by NAR and Tripoli. As we are close to
this with a static stability margin of 2.67 the rocket should be fine. Static stability margin is a
massive part of the failure analysis as it is a failure of the rocket as a whole if it were to head off
in a wrong direction due to instability. It can only be mitigated by properly designing the rocket
to be stable which in this case, the rocket has met the above criteria.
The next point of failure of the rocket will be failure point involving the fins. As the
rocket is moving upwards if the fins are not of the proper shape as higher speeds are reaching the
fins will begin to flutter. In some catastrophic situations the fins can flutter violently enough to
shred and render them and the lower region of the rocket useless. In order to combat this we first
began to look at our fin design from the PDR. We took note of its shape and span. Upon
reviewing the old design and then following it with research we were able to conclude that we
needed to revise the design. We did this by reading NACA Tec Note 4197 and “The Flutter
Problem, Modeling Unstable Aeroelastic Vibrations in Rocket Fins” , a research paper by
Vincent San Miguel who is a Penn State student. Using the equations and knowledge garned in
these papers we were able to combine that with an Excel spreadsheet calculator to produce data
that would model and project fin flutter in our fin design. We found that by shortening the span,
increasing the root chord and changing the shape, that flutter would cease to be a problem. This
test has yielded the results, that are based in data, that by using a clipped Delta Shape, that fin
flutter will not be a problem in this situation. Infact it was found that we had a 20% margin of
safety giving us a 120% stability when faced with fin flutter. This in turn mitigates the point of
failure that is known as fin flutter during flight.
There is also a point of failure in the way the fins are attached to the rocket. Using a
slotted body tube the fins will be mounted to the motor mount of the rocket by a method referred
to simply as “through the wall fins.” The fins will then be filleted by an epoxy putty then further
secured with a 2 inch band of carbon fiber running the length of the root chord. This will not
only mitigate the point of failure that is where the fins attach, but it will also reduce the
coefficient of drag on the rocket.
The point of failure is the airframe. The airframe will be prone to bending and flexing
during the rockets flight. However given that the body tube is composed of fiberglass it is not
likely to bend or flex. This is merely an assumption so in order to properly assess the risk of
failure when it comes to the body tube we did research followed by calculations. Our research
was that of looking into the Contest - Roc Yahoo group and using their Flex- Roc calculator in
order to plug in our rockets components and materials to predict whether or not it will flex and
break during its flight. The program ran many iterations of the formulas involved to predict the
materials failure during flight every step of the way. It was solving for an aeroelastic divergence.
The software found that our rocket was not in fear of aeroelastic divergence as the team had
predicted, and therefore the rocket was not at risk of flexing to a point of destruction. So this
point of failure was properly mitigated by picking a strong enough material to house the rockets
internals for flight.
The next point of failure is the recovery system. will be as follows. To start with the
chute material will be a Rip Stop nylon with a rating of 70DEN FR to be provided by Seattle
Fabrics. The FR designation means fire retardant, as the chute is a polyurethane coated fabric.
With the chute diameter measured out to eighty-four inches we will be using an eighth inch Flat
Braid Dacron which is rated up to four hundred pounds. It can be found at paragear under the
category number W9754FR. We chose the eighth inch Dacron over the three sixteenth inch as it
is lighter and conserves space this way. Also the eighth inch Dacron is strong enough as each
line is rated up to four hundred pounds. So with there being eight gores this gives our shrouds the
cumulative strength of thirty-two hundred pounds, or about 320G's. Meaning it would take an
opening shock at 320G's of force to cause a failure in the chute material or shroud lines, which is
extremely unlikely to happen at it's time of deployment. Also as far as the shroud line goes we
are using a flat fabric as opposed to round due to it being easier to sew with our given skill set.
Two types of thread line will be provided by Seattle Fabrics for use in the chute. The first being
Z46 thread used to sew the gores together. It is size forty-six and one ounce thread that is nylon
bonded giving it the Z designation. The next being the Z69 to sew the gores and threads to the
shroud lines. It is size 69 and nylon bonded giving it the Z designation. As mentioned prior we
plan on using eight gores in the construction of the chute. The gore length is fifty inches for this
semi hemispherical design with a gore width of thirty three inches. It has an apex of forty-five
and base angle of sixty. The base angle is derived from the fact that a semi hemisphere is two
thirds of a full hemisphere leaving us with an angle of sixty instead of ninety. Using loops we
will be able to give each piece two shroud lines sewn the length of the gore giving it reinforced
strength. This also makes it easier to hook the parachute to the hardware as it will be presented in
a loop instead of leaving us with shrouds to knot. We will be running French fell seams along the
entirety of the chute which is a strong stitching pattern. This clearly has been determined to be a
strong pattern for our main chute being the semi hemispherical one. The X-form drogue has
already been sewn in house with all materials and techniques aforementioned. It has so far
successfully brought down a level two rocket flying a J class motor that weighed in at ten pounds
with a height of eighty four inches.
Next we have the hardware. The shroud lines being composed of loops will be attached to
to a D Link. This D link will come from a standard hardware store and is size five sixteenths.
This gives it a strength rating of 1500 pounds or 150G's. The D Link will hook into a barrel
swivel that is provided by LOC precision. It is swivel 12 and rated for 1500 pounds of strength.
Which is again at 150G's of force in order to induce failure. The barrel swivel will be hooked
onto our quarter inch tubular kevlar shock cord which carries a 3000 pound strength rating.
Which comes down to 320G's of force to induce failure. The shock cord will be tied onto a Ubolt that will be anchored in a carbon fiber bulk head made with 24KPL50 which provided by
solarcomposites. According to their website it boasts a rating of aerospace 600KPSI.
Through the parachute design and hardware applications the point of failure is
determined to be the D-link or barrel swivel which are rated for 150G's of force or 1500 pounds.
However seen as at apogee the rocket will not nearly be under this pressure and according to the
rate of descent calculator once the drogue is deployed it will descend at 24 feet per second.
Meaning even when the main chute deploys it will still not experience 150G's of force, thus
giving our chutes a theoretical point of failure at the D-Link to barrel swivel connection however
it is not likely to happen. It has been properly mitigated by analyzing all possible points of failure
then looking at the feasibility of it actually happening.
The next point of failure during the failure analysis is the electronics components. The
electronics have three points of failure. First they can fail due to improper wiring, which would
cause the electronics to malfunction and not work at all. Second they can fail due to an improper
power source, without power they cannot run. Finally they can fail due to manufacturer defect. In
order to mitigate against these points of failure in the electronics, they must be thoroughly
inspected for the faults. Then the appropriate action must follow, whether it is changing batteries
or adjusting wiring. It can be difficult to mitigate for a manufacturer defect however testing upon
receiving the components should work well towards combating it.
Another common point of failure is the nose cone. As the nose cone is held onto the
rocket by a shock cord, special care must be taken to mitigate the possible of a failure of the nose
cone to be held to the rocket. The team is attempting to mitigate this risk by using a nose cone
formed out of a kevlar and fiber glass composite to give it structural rigidity, as well as attaching
it firmly to the shock cord using strong materials such as metal bolts and metal wire. This will do
well to prevent a point of failure in the nose cones attachment to the rocket.
During the failure analysis the boat tail has also come under scrutiny. The boat tail could
present a potential point of failure if it is not firmly attached to the rocket. In order to mitigate
this the boat tail will be attached to the rocket by using an epoxy resin. Epoxy resin should be
strong enough to hold the boat tail to the rocket as the boat tail isnt taking a large force of
pressure during flight.
There is also the thought of the rail guides. In order to assure the rail guides do not shear
off once the rocket is loaded onto the rail, they will be bonded to the airframe with epoxy which
should be more than capable of holding them on there. This will be mitigated by the team by
ensuring the strength of the epoxy once they have been mounted.
The rockets over all safety has been determined by this failure analysis. The safety of the
rocket will be quite assured once the rocket is assembled and ready for flight. Additional safety
hazards the rocket may be concerned with outside of the aforementioned failure points, are it’s
shipping and transportation to the launch site. Transportation is a big logistic of such a large
rocket and as such it is mission critical that no parts be damaged during its transport. This
presents a large concern for the rockets safety however there is not much the team can do to
mitigate this safety concern if the team decides to have the rocket shipped as it is entirely up to
the people who handle the rocket during its transit. Although the team can take preventative steps
to package it to try and reduce and damage due to mishandling during transport.
List of personnel hazards and demonstration research is as follows.
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ROL ATF News
Federal Explosives Law (790k)
Exemption Paper Trail (613k)
List of Explosive Materials (96k)
Federal Firearms Laws (227 k)
Letter from ATF 1-13-99 (57k)
AeroTech to J. Zamillo 11-4-98 (70k)
ATF Communique 4-25-94 (53k)
Memorandum of Opinion 6-24-02 (1.48MB)
Letter From ATF 4-20-94 (80k)
Letter From ATF 6-20-94 (66k)
G. Brunda to ATF 5-28-92 (104k)
ATF to G. Brunda 6-5-92 (21k)
ATF to T. Smith 11-28-01 (35k)
GCR ATF Affidavit 6-27-02 (217k)
Court Order re: TRA/NAR vs. ATFE 3-16-09
Environmental concerns.
Environmental concerns are largely confined to the specifics of the Bonneville Salt Flats.
To begin with possible environmental concerns start with the weather and conditions of the
launch site. Rain is an environmental concern that carries the capability to possibly damage
electronics. There is also that of salt of dust which are concerns for an area as arid as the salt flats
as they can potentially damage electronics through corrosion or interference. The rocket also
faces the environmental concern of the ground. The salt flats have a rough compact surface
which can be potentially damaging to the rocket depending on its rate of descent as it lands on
the surface. Also the rocket carries the potential of landing near hazardous indigenous wildlife
that is native to the environment present. In the event of this, it could also pose a threat to the
team during the retrieval process of the rocket. Also the team faces the environmental concerns
faced with human health in a dry environment. This means proper hydration as well as taking
precautions to make sure no one injures themselves through accidents on the flats. Such as
falling and scraping their knee as an obvious example. That is a viable environmental concern as
rough ground is extremely hazardous towards fragile human skin. Simple mitigations towards
these environmental concerns are as easy as taking the obvious proper precautions such as
bringing water or avoiding the native wildlife.
V) Project Plan
Show status of activities and schedule
Budget plan
Funding plan
Funding for the rocket and travel arrangements will partly be provided by the University
of Central Florida Student Government Association. The remainder of needed funds will be
fundraised by the members of the team. The team first plans to begin an IndieGoGo campaign in
order to raise a large portion of the funds given the success of other SEDS chapters abilities to
raise funds for their rocket projects using the platform. Additionally the team would also like to
reach out to local companies for sponsorship since the team is located just shy of the “space
coast”. With a very space oriented community it should be plausible to locate sponsors and
donors.
Timeline
Educational engagement plan and status
Team Project Advance plans to make an educational presentation at Jackson Middle
School. Following the presentation students will participate in a question and answer session
regarding STEM fields. Additionally, the team will show the students data from previous
launches and projects that members of Project Advance have participated in. The team will also
teach the students how to build paper rockets and the physics behind them. The students will be
able to launch them on a compressed air rocket launcher in order to observe these concepts in
action.
VI) Conclusion
The University of Central Florida is eager to compete once again in the NASA Student
Launch project. The members of Project Advance believe that their payload will be successful
and the data collected from this experiment will be significant. Through testing and
experimentation, the team plans on having a successful building process and launch. The
payload, which includes a live-feed camera, strain gauges, voltmeters, and an accelerometer, will
be crucial to the development of future UCF projects and NASA missions.