Critical Design Review
University of South Alabama Launch Society
Conner Denton, John Faulk, Nghia Huynh,
Kent Lino, Phillip Ruschmyer, Andrew Tindell
Department of Mechanical Engineering
150 Jaguar Drive, Mobile, Al, 36688
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Contents
I
Summary of Critical Design Report
5
I.A Team Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5
I.A.1
Team Name . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5
I.A.2
Mailing Address . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5
I.A.3
Team Mentor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5
I.B Launch Vehicle Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5
I.B.1
Size and Mass . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5
I.B.2
Motor Choice . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5
I.B.3
Rail size . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5
I.B.4
Recovery System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5
I.B.5
Milestone Review Flysheet . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6
I.C Payload Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6
I.C.1
Summary of Payload Experiment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
II Changes Since Preliminary Design Review
6
7
II.A PDR Feedback Questions and Answers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7
II.B Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7
III Vehicle Criteria
8
III.A Mission Statement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
8
III.B Success Criteria . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
8
III.C Major Milestone Schedule . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
8
III.D Vehicle System Level Functional Requirements and Verification . . . . . . . . . . . . . . . . .
9
III.E Workmanship . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
11
III.F Additional Testing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
11
III.GRemaining Manufacturing and Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
12
III.H Design at a System Level . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
12
III.H.1 Drawings and Specifications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
12
III.I Integrity of Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
14
III.I.1 Fin Suitability . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
14
III.I.2 Proper Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
16
III.I.3 Proper Assembly Methods . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
16
III.I.4 Motor Mounting and Retention . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
17
III.I.5 Mass Statement
. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
17
III.I.6 Safety and Failure Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
19
III.J Recovery Subsystem . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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III.J.1 Parachutes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
21
III.J.2 Electrical Hardware . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
22
III.J.3 Kinetic Energy . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
23
III.J.4 Safety and Verification Plan of Recovery System . . . . . . . . . . . . . . . . . . . . .
23
III.J.5 Risks & Mitigation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
25
III.J.6 Recovery System Risks & Mitigation . . . . . . . . . . . . . . . . . . . . . . . . . . . .
26
III.J.7 Payload Risks & Mitigation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27
III.J.8 Booster Bay Risks & Mitigation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
29
III.K Mission Performance Predictions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
30
III.L Drag . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
32
III.L.1 Legend for Drag . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
32
III.L.2 Body Tubes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
33
III.L.3 Interference Drag . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34
III.L.4 Launch Lug Drag . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
35
III.L.5 Stability . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
35
III.L.6 Wind Drift . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
37
III.MPayload Integration
. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
37
III.N Launch Concerns and Operations Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . .
38
III.N.1 Final Assembly Checklist . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
40
III.O Safety and Environment (Vehicle and Payload) . . . . . . . . . . . . . . . . . . . . . . . . . .
40
III.O.1 Safety Officer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
40
III.O.2 Preliminary Hazard Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
41
III.O.3 Personnel Hazards . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
41
III.O.4 Failure Modes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
45
III.O.5 Environmental Hazards . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
49
IV Payload Criteria
51
IV..1
PixyCam . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
51
IV..2
DC Turbine . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
52
IV.A Payload Testing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
54
IV.A.1 PixyCam . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
54
IV.A.2 DC Turbine . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
55
IV.B Uniqueness and Difficulty of Payloads . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
55
IV.B.1 PixyCam . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
55
IV.B.2 DC Turbine . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
56
IV.C Turbine Theory . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
56
IV.D Payload Concept Features and Definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
57
IV.D.1 Creativity and Originality . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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IV.D.2 Suitable Level of Challenge . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
59
IV.E Testing and Design of Payload Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
59
IV.F Science Value . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
59
V Project Plan
59
V.A Budget Plan . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
59
V.A.1 Rocket Structure Budget . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
59
V.A.2 Outreach Budget . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
61
V.A.3 Budget for Trips to Samson, Alabama . . . . . . . . . . . . . . . . . . . . . . . . . . .
62
V.A.4 Huntsville Budget Plan . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
62
V.A.5 Total Project Cost . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
63
V.B Funding Plan . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
63
V.C Project Timeline . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
63
V.C.1 Event and Submission Schedule . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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V.C.2 Gantt Chart . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
66
V.D Educational Outreach . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
66
V.D.1 Purpose . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
66
V.D.2 Status of Outreach . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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V.E Schedule for Outreach Events . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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VI Conclusion
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I.
I.A.
I.A.1.
Summary of Critical Design Report
Team Summary
Team Name
The team’s name is University of South Alabama Launch Society and the vehicle is Montalvo’s Minion.
I.A.2.
Mailing Address
The mailing address for the team is 150 Jaguar Dr., Mobile, AL 36688-6168. The team can be contacted
through the Dean’s Office at (251)460-6168 or via email at usals@southalabama.edu.
I.A.3.
Team Mentor
The University of South Alabama Launch Society’s team mentor is Mr. John Hansel. He is a retired
electrical engineer from the Department of Defense. To this date, he holds level 3 certification in both NAR
and Tripoli rocketry associations and has been experimenting with model rocketry for over 15 years.
I.B.
I.B.1.
Launch Vehicle Summary
Size and Mass
The vehicle is projected to be 94 inches in length and weigh approximately 15.758 pounds. A more detailed
description on the vehicle can be found in the Vehicle Criteria section.
I.B.2.
Motor Choice
The motor selected for the vehicle is a reloadable AeroTech K550W-18.
I.B.3.
Rail size
The launch rail selected is 10 feet in length.
I.B.4.
Recovery System
The recovery system will consist of a dual deployment style of ejection where there will be two parachutes
deployed: a main and a drogue. The drogue parachute will be ejected first via black powder charge once
the rocket attains apogee. As the rocket is descending, the main parachute will be ejected via black powder
charge at a fixed altitude. The use of altimeters is vital in order to execute this method of deployment. The
parachute selection was based upon the drag characteristics in order to obtain a safe, smooth descent for the
recovery of the vehicle.
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I.B.5.
I.C.
Milestone Review Flysheet
Payload Summary
There will be two payloads integrated upon the vehicle: the hazard detection system and the proof of
electricity generation with a wind turbine.
I.C.1.
Summary of Payload Experiment
For the hazard detection payload, a Pixy CMUcam5 sensor will be programmed by an Arduino Uno circuit
board. The Pixy cam will be programmed to detect change in color variations specifically to the user’s
desires. It will also be able to capture images, by command of the user, to record data and images. The
wind turbine payload will be able to generate an electrical current that will be sufficient to power drogue
parachute deployment. Small holes will be drilled into the section preceding the wind turbine, and small
scoops attached to the exterior of the rocket, in order for air circulation to come in. Since it is not within the
competition regulations to power the deployment on the vehicle this way, a chart of the voltage generated
will be produced to prove that this method would work.
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II.
II.A.
Changes Since Preliminary Design Review
PDR Feedback Questions and Answers
Table 1 displays the feedback the team received from the preliminary design report and the solutions provided
for each critique.
Table 1: PDR Critique and Solutions
1
Critique
Your apogee is listed as 6,000 ft. This is well above
our FAA waver limit of 5,600 ft. Please prove to the
review panel that your apogee will not exceed 5,600
ft. by CDR.
2
Report mentions the use of cardboard as a construction material. Can you confirm that this is a phenolic
substance, and not card board?
3
The stability margin of 1.81 is low, and our chief
engineer would like to see a stability of 2.0 as the
rocket leaves the rail.
You have a high thrust to weight ratio. What are
the construction techniques for your fins?
4
5
6
You have a slow descent velocity which can lead to
excessive drift. In general, descent velocities of 80100 fst. and 17-20 fst. for the drogue and main chutes
respectively will work for a rocket of your size and
weight.
What are you generating electric current for?
7
Where are you placing your turbines?
8
Does your turbine use an air scoop? If so, how does
it affect the stability of the rocket?
II.B.
Solution
USA Launch Society has changed the vehicle’s motor
brand from Loki Research to Aerotech. With that
in mind, the team has changed the motor from a
K960SF to a K550W motor, thus results in a lower
simulated apogee of approximately 5,160 feet.
The team has changed the vehicle’s body material
along with the motor tube and centering rings to G10 fiberglass. There will be brief stress and buckling
analysis provided in the vehicle criteria ratio. III.I.5
The team has an updated OpenRocket simulation
that results in a stability margin of 3.6 cal.
Our thrust to weight ratio was initially miscalculated
in the PDR. The current thrust to weight ratio for
the new configuration for the vehicle is 6.15.
The team has taken into consideration the recommended descent velocities. Team members have performed analytical computations and simulations that
results in a descent velocities of 131 fst. and 22.07 fst.
for the drogue and main parachutes, respectively.
The team will be generating electric current to serve
as a theoretical back-up voltage supply in order to
deploy the drogue parachute. Data will be logged on
an Arduino from the turbine to measure the voltage
to ensure that the voltage is high enough to theoretically ignite black powder to deploy the drogue
parachute.
The team has chosen for the turbine(s) to be placed
in between the CP and CG to ensure a better stability of the rocket.
The team has designed air scoops to be placed over
holes machined through the booster bay and turbine
payload. The scoops will be used to direct the air
flow over the turbine. A scaled down version of the
rocket vehicle with scoops will be tested in the wind
tunnel at USA.
Overview
In the aspects of the vehicle and payload criteria, there were not any significant changes made. Parameters
such as the dimensions, material usage, and a more precise mass is defined throughout the report for the
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meager changes. Furthermore, this report will provide a more intricate description of each element that the
vehicle will sustain and more accurate calculations and simulations. Also, the team will provide the material
structure for each discrete element such as G-10 fiberglass and BNC-20B balsa wood bulkheads.
III.
III.A.
Vehicle Criteria
Mission Statement
The stability of the launch vehicle is based upon the intricate design in selecting the size, mass, motor choice,
and recovery system. The team will analyze each subsystem and determine the most appropriate design to
provide the best overall performance of the vehicle in all aspects. Design, review, and simulation of each of
these subsystems will be an important part in the success of a stable vehicle design. As a result, the team
will work to produce a versatile, unique vehicle design that will be a viable competitor against others in the
USLI competition.
III.B.
Success Criteria
As far as the what is required in succeeding in the vehicle aspect of the design, this will consist of creating a
presentable design that will be highly communicable between the team and others. The team will determine
ideas and work towards designs in SOLIDWORKS that will help illustrate the design of each subsystem such
as the overall vehicle, nosecone, parachute bay, payload bay, booster section, and fin design of the rocket.
III.C.
Major Milestone Schedule
The Gantt chart shown in Figure 23 displays the team’s project plan with any and all major milestones and
plans included.
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Figure 1: Gantt chart showing the projected schedule for the team to complete competition tasks.
III.D.
Vehicle System Level Functional Requirements and Verification
Below, Table 2 demonstrates that all of the vehicle’s systems are defined and verified, with reasoning.
1
Vehicle Requirement
Requirement Features
The vehicle should deliver two
In order to achieve this altitude, the aerodynamics of the rocket will be maxi-
payloads to an altitude of 5,280
mized by selection of the ogive nose cone and the straight-tapered fins. Launch-
feet from the ground.
ing with the correct motor size will also greatly contribute to the altitude goal.
In order to ensure that the rocket is capable of flying to 5,280 feet, the team
will use OpenRocket to simulate the launch.
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2
The official altitude will be deter-
USA’s rocket will contain two barometric altimeters. One altimeter will control
mined using one, on-board baro-
both the main and drogue deployments while the other altimeter will serve as
metric altimeter. The rocket will
the official scoring altimeter.
contain one altimeter for both
deployments and one predetermined to be the scoring altimeter.
3
The vehicle must be reusable
The exterior of the launch vehicle will be coated in a selected material to
and recoverable after each launch
increase the strength of the body.
with minimal or no damage.
4
5
The launch vehicle shall have a
USA will have a two piece rocket. One section will carry the drogue deployment
maximum of four independent
and the turbine system while the second section will contain the main parachute
sections.
and the hazard detection camera.
The launch vehicle shall be lim-
USA’s launch vehicle will perform its only motor ignition at takeoff.
ited to a single stage.
6
7
The vehicle must be quickly pre-
Prior to the launch date, the team will practice preparing the rocket to ensure
pared for flight within two hours.
timely delivery.
The launch vehicle shall be capa-
When the team is practicing the timed vehicle assembly, they will also test to
ble of remaining in launch-ready
ensure the rocket maintains functionality on the launch pad for a minimum of
configuration at the pad for a
one hour.
minimum of 1 hour without losing the functionality of any critical on-board component.
8
The launch vehicle shall be capa-
The launch vehicle will be capable of accommodating the 12 volt firing system
ble of being launched by a stan-
by utilizing e-match insertion into the motor.
dard 12 volt direct current firing
system.
9
The launch vehicle shall use a
All USA motors will be purchased from legitimate model rocketry vendors.
commercially available solid mo-
These motors will be NAR, TRA, and CAR certified.
tor propulsion system using ammonium perchlorate composite
propellant.
10
The total impulse provided by a
Currently, USA plans to use a K-class motor which falls short of the 5,120
launch vehicle shall not exceed
Newton-seconds total impulse.
5,120 Newton-seconds (L-class).
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11
Pressure vessels on the vehicle
Since there will be no pressure vessel on-board the launch vehicle, no extra
shall be approved by the RSO.
measures will be taken.
They should have a minimum
factor of safety of 4:1 as well as a
pressure relief valve that sees the
full pressure of the tank.
12
All
teams
shall
successfully
The team plans to launch a subscale of the final rocket by mid-December.
launch and recover a subscale
model of their full-scale rocket
prior to the CDR.
13
All
teams
shall
successfully
launch and recover their full-
The full scale launch will be performed in early to mid-February to ensure
timely delivery of results by the FRR.
scale rocket prior to the FRR in
its final flight configuration.
Table 2: Verification Plan of Launch Vehicle
III.E.
Workmanship
In order to ensure proper functionality of the rocket, the payloads and their components, the team will learn
manufacturing basics from the gentlemen in South Alabama’s machine shop. These two men, John and
Terry, will help to manufacture some components while they will also teach and guide the USALS students
in the processes. It is important to the team to have supervision during some processes to ensure quality
and safety. Simpler tasks such as soldering electrical components, basic drilling, applying adhesives, etc. will
be performed on the team’s own time without mentor guidance. Also, the team has learned how to operate
the university’s 3-D printer and will continue to use it under the guidance of the mechanical engineering
department head, Dr. David Nelson. To ensure that all processes function correctly, the team has developed
attached checklists for safety and quality purposes.
III.F.
Additional Testing
Each subsystem must be fully operational to hold true to the success of the vehicle. First, the team will test
each subsystem individually and make adjustments. Following the individual tests, the subsystems will then
be simultaneously tested to ensure full operation of the vehicle. With this method, each team member can
and will understand the concepts of each subsystem and how each subsystem will respond to one another.
In addition to this, the testing will take place in a safe, open area to evaluate each subsystem at once.
The parachute testing, however, will take place by setting up a deployment contraption with the parachute
packed into a dummy vehicle. The altimeter bay and black powder charge will be installed in the vehicle.
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To test the deployment procedure, the vehicle will be driven up to a certain speed to deploy the parachute.
The payload components will be tested in the university electrical engineering labs as the necessary testing
equipment is located there. Following the testing of the subsystems individually, the team will test the
systems integrated together in the test launches leading up to the competition.
III.G.
Remaining Manufacturing and Assembly
During the preliminary design process, the team worked together to design each component of the vehicle on
SOLIDWORKS to display the integrity of the design. These drafts of each component were vital in ensuring
the tolerances of each component to fit in the vehicle itself. From then, the team began the construction of
each part starting with a master checklist of all the vehicle components. To create the inner components
of the payload bays, such as the altimeter and payload electronics trays, the team used the university 3-D
printer. Once all parts created by the 3-D printer were finished, they were fitted together to test for, and
mitigate, any clearance issues. To maximize stability and reach the goal altitude of the competition, the
team must design the exterior of the vehicle to be aerodynamically sound while creating high resistance to
outside atmospheric conditions. Through down selecting on many designs demonstrated in earlier sections
of the nose cone shape, overall vehicle body, and the shape of the fins, the aerodynamics of the vehicle can
be maximized. A minimal drag coefficient is essential for the nose cone as it is the forefront component of
the rocket. The ogive shaped nose cone proved to satisfy this requirement. The four fins work to bring the
most stable flight while only minimally increasing drag and using the straight tapered designed fins will also
minimize drag. The inner material of the vehicle will be a selected material by the team to increase strength
and durability of the rocket. The final finish of the vehicle will be chemically coated to produce the least
air friction. The team has begun designing the scoops that will allow air to flow into the rocket and power
the turbine in order to generate current. The flight data and load analysis is yet to be performed on these
scoops.
III.H.
Design at a System Level
III.H.1.
Drawings and Specifications
The critical design of the launch vehicle will compartmentalized into 3 sections: the nosecone, P.A.P.
(parachute, altimeter, payload) bay, and the booster bay. The nosecone will be the leading section of
the rocket that will be of main importance of the rocket’s flight through the atmosphere and the recovery
stages of flight. The P.A.P bay will contain the vital electronic components such as the necessary altimeters
and hazard detection payload components of the finalized rocket. The booster bay will be the trailing section
of the rocket which will entail the fins, finalized motor and its components, and the turbine current generation payload and its components. A detailed description of the vehicle sections that involve the nosecone,
body-tube, fins, motor mount, altimeter tray section of the P.A.P bay will follow.
The rocket body-tube and fins of the final vehicle will be fabricated by G-10 fiberglass. This material was
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downselected due to its high strength and ability to withhold its stability over critical temperatures. This
material is also obtainable by the team. The nosecone will be made of PVC-like material that will have
adequate aerodynamic properties and durable strength. The motor mount and centering rings of the rocket
will be comprised of G-10 fiberglass while the bulkhead will be made of balsa wood. These materials will be
used for its cost effective properties and its light mass. These materials are relatively high in strength for
the stress applied on the material.
The final fin design will be a stream-line straight tapered on each of the four fins. This fin design is most
sufficient for a low drag coefficient with the vehicle traveling at such speeds of subsonic and trans-sonic
flights. Also, this design is more efficient to construct from the material used. The taper ratio that the
team will use will be of 0.52, with a tip chord of 3.25”, and root chord 6.25”. More details on the process
of the fin design is discussed below. A rocket’s nose cone is a vital aspect of the vehicle being that it is
the foremost component and its use throughout flight. The determination of each nose cone in any rocket
comes with heavy considerations of aerodynamics. The main variable, however, to consider for a nose cone
is the drag characteristics it will employ in flight. The South Alabama team wanted to consider using a nose
cone shape that will consist of minimal drag coefficient, yet is durable in its shape and material. The final
nosecone shape that the team carefully chose was an ogive shape made with PVC-like material. This shape
gives very minimal drag characteristics and performs very well in the mach number that will be experienced
during flight. These characteristics of the rocket are shown in the final design and specifications below.
Figure 2: Final rocket drawing.
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Figure 3: Final rocket dimensions.
III.I.
III.I.1.
Integrity of Design
Fin Suitability
The purpose of putting fins on a rocket is to provide stability during flight, that is, to allow the rocket to
maintain its orientation and intended flight path.7 Fins provide a high restoring lift force at even small angles
of attack; therefore, it reduces the turning momentum of the rocket drastically while keeping it dynamically
stable. Fin selection is typically distinguished into two factors: shape and number of fins. The number
of fins considered by the team were either 3 or 4. Equation (1) will be considered in accordance with the
determination of the number of fins.
Fdrag =
1 2
ρV As CD
2
(1)
The team finalized that 4 fins would add more weight; however, the extra stability is more critical and
beneficial in the long run although it will add slightly more drag than three fins would. Along with the
number of fins, the shape of the fins is arguably the most crucial factor. The reason for it being crucial is
because different shapes and sizes will give discrete drag characteristics. After researching different shapes
and airfoils, the team has chosen to use a streamlined straight-tapered fin. First of all, the team analyzed
three designs for the lead and trailing edge of each fin. These airfoil designs are square, rounded, and
streamlined. These designs can be seen in Figure 4 where the top airfoil is square, the middle airfoil is
rounded, and the bottom airfoil is streamlined. Using the same figure, it can be seen that as velocity
increases, the drag rises exponentially with no regard to airfoil design. However, this exponential growth
due to velocity is less rapid for the streamlined airfoil compared to its counterparts.
As previously mentioned, the team has decided to use the straight-tapered planform. Four planforms
were considered for the fin design. These planforms are shown in Figure 5. The team chose not to use
the rectangular planform since it has the highest surface area, which increases drag. Generally, the swepttapered planform is used when rockets will be traveling at supersonic speeds, due to its ability to break the
shock-wave barrier produced at this velocity. Since the competition rocket will be traveling subsonically, the
swept-tapered is not necessary. Elliptical fins are a far less common planform. Due to the complexity of
constructing this planform, the team chose to remove it. Now it is evident that the straight-tapered planform
is the best option because it is easy to construct, it is generally used it subsonic to trans-sonic flight, and it
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has a reduced surface area compared with the rectangular design.
Figure 4: Relation of velocity to drag due to the airfoil shape of the fin.
Figure 5: Common model rocketry planforms.
An important aspect of the straight-tapered fin design is the taper ratio, λ. This relation is shown in
equation (2). OpenRocket software will be utilized to produce tentative fin dimensions. Currently, the South
Alabama Launch Society is using a taper ratio of 0.52, where the tip chord, CT , is 3.25” and the root chord,
CR , is 6.25”.
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λ=
III.I.2.
CT
CR
(2)
Proper Materials
The structural elements of the rocket body and fins will be composed of G-10 fiberglass. Structural elements
include, but not limited, to centering rings, couplers, and electronic bays. The determination for G-10
fiberglass is based on the availability from our supplier and the successful results of the stress analysis that
will prove the material will be able to withstand the loads. The bulkheads will consist of BNC-20B balsa
wood for its light weight, tested durability, and availability to the team. The nosecone will be made of
material similar to PVC because of its appropriate density characteristics and the successful consistency
from past test flights.
III.I.3.
Proper Assembly Methods
The materials used in the vehicle assembly will consist of pre-manufactured body tubes, fins, couplers, and
nosecone will be bought and cut according to the design specifications and dimensions as seen in section
III.H. Manufactured cardboard and balsa wood will be bought and cut by Team South Alabama according
to the design specs. A 3D printer was used to create the altimeter trays and scoops as used in the vehicle
design. These components will be primarily connected through the use of epoxy. The use of epoxy is a
common practice in amateur rocketry.
The hazard detection payload will be placed approximately 36 to 40 in. from the bottom of the rocket
inside the booster bay. A black powder charge will be set to deploy the droque parachute inside the P.A.P.
bay, and the hazard detection payload will be connected to the shock cord. This connection will allow the
hazard detection payload to be pulled out of the booster bay, and thus suspended from the rocket vehicle
to allow full surveillance of the ground surface below when descending. The turbine payload will be place
between the center of pressure and center of gravity at approximately 24 in. from the bottom of the rocket
inside the booster bay to improve stability of the vehicle. Holes will be drilled through the booster bay and
the turbine payload, and 3-D manufactured scoops will be placed over the holes along the outside of the
booster bay in order to direct air-flow through the turbine payload III.H. A coupler will be used to connect
the P.A.P. bay to the booster bay. The Figure below illustrates the placement of all vehicle components
inside the rocket body as designed. A label is given either above or below each component, as seen in the
Figure.
Figure 6: Internal Components of Rocket Vehicle.
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The parachute will be packed as taught by rocket expert Kendall Brent from the Samson launch site
where team members earn their certifications. The motor packing technique will be based off of directions
from the manufacturer. Along with the manufacturer directions, Mr. Brent will shadow the team to ensure
safety and the appropriate protocols are abided.
III.I.4.
Motor Mounting and Retention
The motor shall be mounted in a fiberglass motor tube that will satisfactorily withstand the impulse of
an Aerotech K550W motor. The motor mount will be supported by four fiberglass centering rings evenly
distributed along the tube. For motor retention, a 2.13 inch (54mm) Aeropack motor retainer will be used
to prevent the motor casing from falling out of the rocket. The motor retainer will look similar to the one
shown in Figure 7.
Figure 7: Aeropack motor retainer.
III.I.5.
Mass Statement
Table 3 shows the current estimated mass of the rocket. These predictions are made by actually weighing
many of the individual components on a digital scale and estimating a few of the more obscure components
such as the turbine payload. Two components with uncertain dimensions at the moment, were weighed
based on some of USA’s components from the past. These include the payload trays and the gyroscope.
Also, the fins, centering rings, and bulkhead masses were calculated by determining the proper volume and
using the respective densities for each material, since
m=Vρ
(3)
According to Table 3, the mass of the rocket on the launch pad with full propellant is 15.758 lb. However,
after the flight, once all the propellant has been burned and the black powder charges have been ignited,
the rocket mass will become 14.44 lb. It should be noted that this final mass differs from the ”Total Mass
(Empty Motor)” mass shown in Table 3 because the final mass mentioned above subtracts the propellant
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weight while the ”Total Mass (Empty Motor)” still has the propellant weight included.
The mass estimate in this report is only a preliminary quote. The final design will possibly have a
different mass for a number of reasons. A primary source of error is the mass of the turbine system. Since
this payload is in the early stages of development, it is very difficult to get an accurate estimate of its mass.
With this knowledge, the team is budgeting up to a 20% mass increase.
After calculations in OpenRocket, the team has calculated that adding 4.56 lb to the rocket mass of
15.758 lb, will make the rocket fly to an altitude of 5,000 feet. For every one foot that the rocket flies under
the goal altitude of 5,280 feet, one point is deducted. For every foot over 5,280 feet, two points are deducted.
Since the competition devotes one hundred points to the the altitude of the flight, realistically, the rocket
cannot fly over 5,380 feet or under 5,180 feet. However, the team has decided that the rocket has a margin
of mass error of 4.56 lb. This will make the rocket fly to an altitude of 5,000 feet. The team shot low on
this mass error to try to reduce the amount of mass added in the future.
Component
Mass (lbm)
Nosecone
0.660
Body Tube (2)
1.920
Motor Mount
0.302
Fins (4)
0.952
Motor (empty)
1.318
Motor (loaded)
3.278
Motor Case
0.507
Motor Retainer
0.090
Body Tube Coupler
0.150
Turbine Bay Tube
0.250
Electronics Bay
0.200
Camera Bay
0.200
Turbine Prediction
0.580
Centering Rings (4)
0.168
Bulkheads (7)
0.770
Main Parachute
0.4425
Drogue Parachute
0.068
Shock Cord
0.9425
Black Powder
0.010
Cellulose Insulation
0.150
Wadding Cloth
0.650
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Rods (4)
0.450
Gyroscope
0.22
Payload Tray (3)
0.300
Arduino Uno
0.060
StratoLogger (2)
0.060
TeleMega
0.060
Battery Pack (2)
0.100
9V Battery (2)
0.200
AT-2B Transmitter
0.060
PixyCam
0.100
Electronics Mounting Hardware (4)
0.050
Eye Bolt With Nut (5)
0.200
Rail Button and Screw (2)
0.020
Nylon Shear Pins (4)
0.020
Total Mass (Empty Motor)
14.44
Total Mass (Loaded Motor)
15.758
Table 3: Mass table of rocket components
The critical design report mass statement in comparison from the team’s previous preliminary design
report has been updated with the consideration of the few changes taken place. One critical change for
the project is the different motor brand and projected level of total impulse. The new motor, an Aerotech
K550W, has a lighter total and propellant weight while still able to reach the desired apogee. The other
masses that have been altered are the shock cord and main parachutes, which are updated to be more precise.
The overall mass was not changed significantly due to the alterations of the aforementioned components being
able to balance out.
III.I.6.
Safety and Failure Analysis
The team is involved with plans to evaluate and verify the vehicle’s stability and performance through
analyzing the vehicle criteria that is set by NASA and making changes to each vehicle subsystem to fit the
requirements. Once the subsystems have passed the vehicle verification, the team will go into further detail
by testing each subsystem to success. A table has been created to outline each vehicle requirement and its
assessment to the design 2.
The stability and drag assessment is provided in the following sections; moreover, a buckling analysis
has been performed to ensure the safety of Team South Alabama’s rocket vehicle. As aforementioned in
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this section, the rocket body tubes will be made out of pre-manufactured G-10 Fiberglass. The material
properties of this material are given in Table 4.
Table 4: Material properties for G10 fiberglass.
Mechanical Property
Fracture Strength
Flexural Modulus
Poisson’s Ratio
Compressive Strength
Value
38000 psi
2700000 psi
0.12
65000 psi
These material properties were originally declared by the ASTM D790 for the properties of reinforced
plastic and electrical insulating materials and retrieved from MatWeb, an online material property database.9
Using these material properties and the vehicle dimensions, hand calculations could be performed to determine values for critical stress and critical load. Since the length of the body tubes is much greater than the
diameter, the buckling analysis was conducted for long cylindrical shells. The equation used to determine
the critical stress is given below:15
Eh
σcr = p
a 3(1 − ν 2 )
(4)
Inserting the appropriate values into the above equation the value for critical stress could be determined,
where a is the radius, h is the thickness, ν is Poisson’s ratio, and E is the flexural modulus. The values
determined from this buckling analysis are as provided in the following table:
Table 5: Buckling calculation variables.
Property
Length (l)
Radius (a)
Thickness (h)
Critical stress (σcr )
Critical load (Fcr )
III.J.
Value
94 inches
2 inches
0.0625 inches
49068.5 psi
3904.75 lb.-f
Recovery Subsystem
The recovery subsystem is a combination of a few different components. There will be two altimeters aboard
the rocket: one for the drogue and main deployments and one for the official scoring. Also, there will be
a transmitter to relay the rocket’s location in case it flies to an unseen location. These electronics will be
attached to igniters which, in turn, are attached to the black powder charges. These charges will be used to
blast the drogue and main parachutes, as well as the hazard detection camera, out of the rocket.
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III.J.1.
Parachutes
The parachute that will be used for the drogue deployment will be solid-bodied with a 12 inch diameter.
The main deployment will be a 72 inch diameter, solid-bodied parachute. The drogue will be attached at the
aft end of P.A.P. bay by six drawstrings. These strings will attach to an aluminum carabiner, which in turn
attaches to a tied loop in the shock cord. This same procedure is repeated for the main parachute, however,
the main will be attached to the shock cord and nosecone at the forward end of the P.A.P. bay. The shock
cord to be used is half inch tubular nylon. The drogue and main parachutes can be seen in Figures 8 and 9,
respectively.
Figure 8: The determined 12” drogue parachute downselected by the team.
Figure 9: The determined 72” main parachute downselected by the team.
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III.J.2.
Electrical Hardware
Figure 10: PerfectFlite Stratologger altimeter that will be used for both drogue and main parachute
deployment phases.
As shown in Figure 10,a StratoLogger altimeter will be used to perform the parachute deployments. This
altimeter will be attached to wires connecting the altimeter to the igniters. The altimeter will be placed in
an altimeter bay with bulkheads on either end. The igniters will be mounted on the exterior of the bulkheads
facing both forward and aft. The wires connecting the altimeter will be attached to the igniters through
very small holes in the bulkheads. Then, the igniters will be buried into the black powder charges. The
drogue parachute will be ignited at apogee, determined by the lack of positive pressure change. Then, as the
rocket descends back down towards the earth, the main parachute will deploy at 800 feet. Also, as previously
mentioned, there will be a transmitter placed in the nosecone of the rocket in order to properly track the
rocket in case the visual of the rocket is lost. Figure 11 shows the electrical schematic of the altimeter hooked
to the main and drogue parachutes.
Figure 11: Electrical schematic of the drogue and main parachutes attached to the StratoLogger altimeter.
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III.J.3.
Kinetic Energy
The definition of a safe landing velocity being that each individual component must not possess more than
75 f t − lbf of kinetic energy at landing. This means that in order to calculate the safe landing velocity we
must first account for the mass of the vehicle and divide it into its respective sections as seen in 15.
Since the Booster bay section has the highest mass, it will also have the highest kinetic energy at any
given speed. The impact velocity of each component when landing is assumed to be the same, and is found
using OPENROCKET to be 20.2 f t/s:
Ek =
1
mV 2
2
(5)
By inserting the determined values, the kinetic energy of each main component can be found using the
above equation. Using the equation for kinetic energy, and substituting the mass of each section and the
ground impact velocity of 20.2 f t/s, the kinetic energies are as follows: Nosecone = 2.79 f t − lbf , P.A.P. =
36.26 f t − lbf , and the Booster = 52.71 f t − lbf .
To determine the kinetic energy of each component at the start of the launch sequence coming off the
rail, the rail exit velocity of 44.2 f t/s given from OPENROCKET is used. Inserting this value of velocity
and the provided mass of each section into the kinetic energy equation, the kinetic energies are as follows:
Nosecone = 13.36 f t − lbf , P.A.P. = 173.62 f t − lbf , and the Booster = 252.37 f t − lbf .
The kinetic energy of each component at the deployment of the main parachute is determined using the
velocity at deployment of 131 f t/s given from OPENROCKET, and the appropriate mass of each component.
Inserting these values into the kinetic energy equation provides the following kinetic energies of each main
component: Nosecone = 117.34 f t − lbf , P.A.P. = 1525.13 f t − lbf , and the Booster = 2216.34 f t − lbf .
III.J.4.
1
Safety and Verification Plan of Recovery System
Recovery System Requirement
Requirement Features
The launch vehicle shall stage
The launch vehicle will use an altimeter at apogee to deploy the drogue
the deployment of its recov-
parachute while a separate altimeter will be used to deploy the main parachute.
ery devices, where a drogue
parachute is deployed at apogee
and a main parachute is deployed
at a much lower altitude.
2
Teams must perform a success-
These parachute testings will be done at the team’s regular launch field prior
ful ground ejection test for both
to both of the aforementioned flights in order to prove recovery system effec-
the drogue and main parachutes
tiveness.
prior to the initial subscale and
full-scale launches.
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3
At landing, each independent
In order to ensure that the independent sections do not exceed the maximum
section of the launch vehicle shall
kinetic energy, the team will model the descent of the rocket in OpenRocket
have a maximum kinetic energy
where the velocity of the rocket at impact with the ground will be calculated.
of 75 ft-lbf.
Then, the kinetic energy of each section will be calculated using the equation
E = 12 mV 2 .
4
The recovery system electrical
Each parachute deployment will be controlled by its own altimeter. Each pay-
circuits shall be completely inde-
load will be controlled by electrical systems that are unrelated to the recovery
pendent of any payload electrical
system altimeters.
circuits.
5
The recovery system shall con-
The launch vehicle will contain two PerfectFlite StratoLoggers and one Tele-
tain redundant,
Mega, all of which are commercially available.
commercially
available altimeters.
6
Motor ejection is not a permissi-
Each StratoLogger will initiate a black powder ejection charge for their respec-
ble form of primary or secondary
tive parachute deployments at apogee and low-altitude main deployment.
deployment. An electronic form
of ejection must be used for deployment purposes.
7
A dedicated arming switch shall
USA will use either simple wire twisting switches or small flip switches to
arm each altimeter, which is ac-
initiate the altimeters from the exterior of the rocket.
cessible from the exterior of the
rocket airframe when the rocket
is in the launch configuration on
the launch pad.
8
Each altimeter shall have a ded-
Each altimeter or GPS will have its own 9 volt battery.
icated power supply.
9
10
Each arming switch shall be ca-
If twisting wires is chosen, they will be tightly secured in a specific hooked
pable of being locked in the
formation. If the flip switches are used, they will automatically be locked into
”ON” position for launch.
position once they are activated.
Removable shear pins shall be
Nylon shear pins will be used to attach the components. The exact properties
used for both the main parachute
of these shear pins will be determined by the CDR to ensure proper shear
compartment and the drogue
calculations.
parachute compartment.
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11
An electronic tracking device
The exact tracking device is yet to be determined at this time. USA’s launch
shall be installed in the launch
vehicle will remain entirely tethered together, therefore, there is no need for
vehicle and shall transmit the
multiple transmission devices.
position of the tethered vehicle
or any independent section to a
ground receiver.
12
The recovery system electronics
The two recovery altimeters will remain in an independent bay from the elec-
shall not be adversely affected by
tronic tracking device in order to avoid inadvertent excitation of the recovery
any other on-board electronic de-
system.
vices during flight.
Table 6: Verification Plan of Recovery System
III.J.5.
Risks & Mitigation
Tables will be used to describe the potential risks of each component of the vehicle and its mitigation
techniques for visual purposes. The sections that will be broken down are as follows: vehicle risks, recovery
system risks, payload risks, and booster bay risks. The NASA Student Launch 2016 Handbook will be
heavily relied on for structuring the following tables.1
Table 7: Potential risks relating to the vehicle as a whole with internal components considered.
1. Risk 1: The vehicle is damaged when ejection charges are ignited.
• Impact: The damage to the vehicle may lead to modifications after landing or failure to reuse the
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vehicle if damage is too extreme.
• Mitigation : Testing of recovery system, specifically with the ejection charges. Weighing the
amount of black powder precisely to ensure the vehicle P.A.P. bay and nose cone do not incur any
damage upon ignition.
2. Risk 2: The vehicle incurs an extensive amount of damage are upon landing impact.
• Impact: The vehicle will not be reusable and the team will receive low scoring in the USLI
competition.
• Mitigation: Proper testing of the recovery system should ensure the vehicle’s landing speed is low
enough to reduce the possibility of any damage.
3. Risk 3: A team member is injured during testing of component systems.
• Impact: Depending on the magnitude of the injury, workload may increase on the remaining team
members.
• Mitigation: Pre-test briefing of all protocol must be performed so all team members are aware of
all possible risks involved and what to do to ensure safe testing.
4. Risk 4: Team members leave the state for school holidays.
• Impact: The team will be prohibited to meet to fulfill project milestones.
• Mitigation: The team has created an Overleaf file which will allow all team members to have
access to write and compile any scholarly documents related to the project.
5. Risk 5: The team may not receive enough funding to purchase all vehicle components.
• Impact: Without sufficient funding, the key components of the vehicle’s design, manufacturing,
and testing and possibly subsystems can be delayed when trying to reach benchmarks.
• Mitigation: The team will reach out to all available sources.
III.J.6.
Recovery System Risks & Mitigation
1. Risk 1: Original planned tests are not successful.
• Impact: Verification of the system is delayed, pushing potential manufacturing dates behind.
• Mitigation : Researching and gaining an understanding of how the subsystems work together to
create the system, to devise a correct test method.
2. Risk 2: System components (altimeters and parachute components) are damaged and/or not functioning as specified.
• Impact: Delays testing and verification of the system and in turn delaying the manufacturing of
the overall vehicle.
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Table 8: Potential risks relating to the recovery system of the vehicle.
• Mitigation: Pre-test briefly specifically covering the handling of the sensitive components. Also,
ordering excess components as a safety net since there tends to be error.
3. Risk 3: Parachute materials are not readily available and need to be ordered.
• Impact: May delay testing plans and/or manufacturing process.
• Mitigation: The team will buy excess amounts of material needed to manufacture the parachutes
in order to hinder the risk from happening.
4. Risk 4: The down selected altimeters are on back order.
• Impact: May delay testing of the electronics, which will create a setback in the team’s specific
benchmarks.
• Mitigation: The team may order multiple of the components or if too close to very important
milestones, the team may select another altimeter that was considered.
5. Risk 5: The altimeters are damaged upon delivery due to poor handling and/or poor packaging.
• Impact: Testing of the recovery system will be delayed in turn pushing important benchmarks
further behind.
• Mitigation: An alternate altimeter that was originally considered will be used.
III.J.7.
Payload Risks & Mitigation
1. Risk 1: Team member injured during testing.
• Impact: Depending on injury, workload may increase on remaining team members.
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Table 9: Potential risks relating to the payloads for the competition.
• Mitigation : Pre-test briefing of all protocol must be performed so all team members are aware
of all possible risks involved and what to do to ensure safe testing.
2. Risk 2: Original planned tests are not successful.
• Impact: Verification of the system is delayed, pushing potential manufacturing dates behind.
• Mitigation: Adequate preparation for planned tests should be done. Research proven testing
methods to reduce the options available.
3. Risk 3: Sensors failing due to too much power.
• Impact: Could melt or damage circuit boards. Delay further testing and increase costs.
• Mitigation: Using test data to verify the amount of input power produces the correct amount of
output.
4. Risk 4: Ordered parts are damaged or broken.
• Impact: Delayed tests for the payload.
• Mitigation: Using reputable sources along with having a supplier to contact in the event that a
component is needed for replacement.
5. Risk 5: Components are damaged or lost during travel to launches.
• Impact: Prevents data accumulation and accurate testing. May require additional travel expenses
to achieve test results.
• Mitigation: Properly package and seal sensitive components. Plan for both loss and acquire extra
parts when applicable.
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III.J.8.
Booster Bay Risks & Mitigation
Table 10: Potential risks relating to the booster bay and its internal components.
1. Risk 1: The motor bay is damaged when ejection charges are ignited.
• Impact: The damage to the vehicle may be amplified upon landing, deeming the vehicle to be
disposable and un-recoverable. Ultimately, the team will fail in the competition.
• Mitigation : Testing of recovery system, specifically with the ejection charges. Weighing the
amount of black powder precisely to ensure the booster bay does not incur any damage upon
ignition.
2. Risk 2: The vehicle does not reach the required altitude for the competition.
• Impact: The team will receive lower scores if the vehicle does not reach or exceeds the expected
altitude originally proposed.
• Mitigation: Communication amongst the team members of any changes made regarding additional
mass and/or dimension changes.
3. Risk 3: Motor is compromised due to moisture.
• Impact: Motor fails to ignite, resulting in a failed launch. Tests can not be performed and data
cannot be collected.
• Mitigation: Properly store and transport motors. Avoid exposing the motor to wet environments
such as rain.
4. Risk 4: Motor fails and/or explodes.
• Impact: Launch is considered a failure. Part of all of the team’s rocket may be compromised.
Sensitive and costly electronics are disposed.
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• Mitigation: Choosing to pack a motor with a kit rather than buy a prepacked motor. Have a
reliable vendor that will sell the highest quality motors in order to reduce the chance of motor
failure.
III.K.
Mission Performance Predictions
USA Launch Society’s mission performance criteria is to achieve the following events:
• Achieve a maximum altitude of 5,280 feet (1 mile), or as close as possible. Note: A maximum altitude
of 5,280 feet is chosen rather than a minimum since the flight score penalty for overshooting the target
altitude is twice as much as undershooting.
• Successfully eject the top body tube and deploy the drogue parachute at apogee.
• Successfully eject the nose cone and deploy the main parachute within the range of 900 to 700 feet as
the rocket is descending.
• Achieve a safe landing and recovery of the rocket with intact parts to be reused.
Figure 12: Flight simulation from OpenRocket that considers no wind speed.
As seen in the flight profile simulations, the predicted point of apogee is 5177 ft. as modeled with a wind
speed of 4.47 mph, and 5182 ft. with wind speed neglected. The maximum acceleration of both simulations
was 377
ft
s2
The simulated vehicle data is as given in the figure:
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Figure 13: Flight simulation from OpenRocket that considers minimal amounts of wind speed.
Figure 14: Simulated vehicle data from OpenRocket.
The weight of each component as well as the center of gravity (CG) and center of pressure (CP) are given
in the following figure:
Figure 15: Analysis of each component of the vehicle via OpenRocket.
The thrust curve given in the figure above, as obtained from THRUSTCURVE.ORG, can be used to
determine the force of the thrust, Fthrust . This thrust curve can be analyzed in order to yield a forcing
function to be used in Newton’s second law, and can also be evaluated to yield natural frequencies for
vibrational analysis. Using the given thrust of 396.8 N from the AeroTech K550W-18 motor and the total
weight of the rocket vehicle being 64.5212 N , the calculated thrust-to-weight ratio is determined to be 6.15.
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Figure 16: Aerotech K550W-18 motor thrust curve for a generic launch.?
III.L.
III.L.1.
Drag
Legend for Drag
ρ = air density
V = velocity
A = cross sectional area of the body tube
CD = drag coefficient
CDN = drag coefficient of the nose shape
CDBT = drag coefficient of the body tubes
CDB = drag coefficient of the base
CDF = drag coefficient of the fins
CDint = interference drag coefficient
CDLL drag coefficient of the launch lugs
(L/d) = total length of the rocket divided by the outside diameter of the rocket
SW = total surface area of nose cone and body tube
SBT = cross-sectional area of nose cone and body tube
Cf = skin friction coefficient
t = maximum thickness of each fin
c = fin root chord
SF = fin surface area
NF = number of fins
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d = maximum diameter of rocket body
CR = the fin root chord length
SLL = cross-sectional area of the launch lug
α = angle of attack
∆CDα = incremental drag coefficient
Drag is a very significant factor in determining the rocket’s performance, and verifying that the rocket is
robust enough to withstand the expected loads. The influence of size, speed, shape, density of the air and
angle will contribute to the drag of the body. These factors are expressed in the basic equation for drag:8
Fdrag =
1
CD ρV 2 A
2
(6)
It is particularly useful to determine the drag coefficient term CD in the equation. Since the drag is
dependent on velocity and air density, noting that these values can change at any instance during rocket flight,
it is more useful to obtain the overall drag coefficient to be multiplied by the other factors in the equation.
Therefore, the following subsections will explain the methods for determining the drag coefficients for each
component of the rocket. The overall drag coefficient for a model rocket is expressed as the summation of
each components as follows:8
CDO = CDN + CDBT + CDB + CDF + CDint + CDLL
(7)
The subscript, O, in CDO is used to represent the drag of the rocket when it is moving directly into the
wind. This is also known as having a zero angle of attack relative to the wind. This will give the lowest
possible value for the drag coefficient, thus resulting in the lowest possible value for the drag of the rocket. It
is good to get the minimum drag first since this gives a goal to try to attain. However, since rockets operate
at various angles of attack, it is important to determine the impact this produces on the rocket’s drag.8 The
current analysis will consist of focusing on determining the drag coefficients at zero angle of attack. This
will serve as the starting point for obtaining an overall drag coefficient. Once a sufficient value has been
obtained for the zero angle of attack case, various angles of attack will then be incorporated to result in a
better approximation.
III.L.2.
Body Tubes
Extensive tests have been performed through the study of aerodynamics to provide a useful formula to
determine the drag coefficient of the body tube with the nose cone included:8
CDN + CDBT
= 1.02Cf 1 +
SW
1.5
(L/d)1.5 SBT
(8)
The skin friction Cf is dependent on the Reynolds number.The Reynolds number will determine if the
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air boundary layer flowing over the rockets surface is laminar or turbulent. For air flows, when the value for
Reynolds number is determined to be less than 100,000, the boundary layer is considered laminar. When
the value is determined to be greater than or equal to 1,000,000, the value is called the critical Reynolds
number (Recr ) and the boundary layer is considered to be turbulent. If the value is in between these limits
(100,000 and 1,000,000), it becomes difficult to accurately predict if the flow is laminar or turbulent. This
is called the transition zone.8 To determine the Reynolds number, the following equation can be used:
Re =
ρV l
µ
(9)
Once the Reynolds number is determined, three approximate skin friction coefficients will be obtained for
three different cases. The first case will be for a fully laminar flow. The second case will be for a boundary
layer that begins laminar and then transitions into a turbulent boundary layer once the critical Reynolds
number is reached. The third case will be for a boundary layer that is fully turbulent. The average skin
friction coefficient equation that will be used for fully laminar flow is defined as:14
1.33
Cf = √
Re
(10)
The average skin friction coefficient equation that will be used for the laminar flow transitioning into
turbulent flow is as follows:5
Cftran =
0.523
1520
−
[ln(0.06Re)]2
Re
(11)
The average skin friction coefficient equation that will be used for the fully turbulent flow is given as:14
Cf =
0.074
Re0.2
(12)
Another significant location on the body tube to consider is at the base (rear of rocket). The drag at the
base is due to the pressure unbalance from the separation of the air flow at the end of the rocket. The base
drag coefficient can be approximated by using the following equation:8
CDB = p
III.L.3.
0.029
CDN + CDBT
(13)
Interference Drag
The interference drag is the result of the air passing over the body tube interacting with the air flowing over
the fins. Sharp corners where the fins meet will increase interference causing a higher drag while smooth,
filleted corners decrease interference causing lower drag.8 The interference drag coefficient can be determine
from the following equation:
CDint = CDOF
CR d
xnumberof f ins
SBT 2
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(14)
III.L.4.
Launch Lug Drag
A method that can be used to determine a value for launch lug drag is to obtain an upper limit on the drag
coefficient. This is done by establishing the worse case scenario for a launch lug. For example, if cylindrical
launch lugs are used, the worse case would be to have a solid disc standing at a right angle to the flow. The
drag coefficient for a disc at right angles to the flow is around 1.2.8 By using the following equation, the
upper limit launch lug coefficient can be determined.
CDLLup = 1.2
SLL
SBT
(15)
Since the launch lugs for this example are cylindrical and not solid flat discs, they will have two surface
areas that experience skin friction drags. The surface area for the cylinder is composed of the inner and
outer surfaces.
SLLw = πdout l + πdin l
(16)
The surface area, SLLw , will then be incorporated into the upper limit launch lug coefficient equation to
obtain the equation for the upper limit coefficient of launch lugs which is written as the follows:8
CDLL =
1.2SLL + SLL
SBT
(17)
This method can also be performed for non cylindrical shaped launch lugs by using the cross-sectional
area and surface areas associated with the different shape.
III.L.5.
Stability
Stability is a significant factor that needs to be considered when designing rockets. It is a must that a rocket
be stable in order to achieve the best and safest flight possible. The first rule to be followed in achieving
stability is that the rocket’s center of pressure must be aft its center of gravity. The distance between the
center of gravity and center of pressure is the static margin, also known as the ”caliber of stability.” The
center of gravity should be at least one body tube diameter fore the center of pressure.6
In 1966, James S. Barrowman and Judith A. Barrowman presented their research and development
project titled, The Theoretical Prediction of the Center of Pressure. The authors successfully derived and
simplified the resulting subsonic theoretical center of pressure equations for general rocket configurations.
The equations found in Barrowmans’ project has provided an accurate method for calculating the center of
pressure for subsonic model rockets and became the staple method for model rocketeers. In recent years,
these equations have been integrated into computer software programs such as OPENROCKET to allow a
quick and easy center of pressure calculation. As the designer changes rocket configurations, the computer
software simultaneously computes the center of pressure and stability margin. Therefore, to simplify the
center of pressure calculation, USA Launch Society will be using the computer program OPENROCKET.
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The computer software approximately located the center of gravity as 59.955 in. aft the tip of the nose
cone and the center of pressure as 74.464 in. aft the tip of the nose cone. This resulting in a stability margin
of 3.6 cal. The center of gravity and center of pressure locations, as well as the stability margin is represented
in Figure 18.
Figure 17: Relationship of the center of gravity relative to center of pressure from OpenRocket
When a rocket is in flight it experiences forces that will cause the rocket to rotate about its center of
gravity. This is called the moment acting on the rocket and is divided into three components: the yawing
moment, the rolling moment, and the pitching moment. These moments are illustrated in the figure below.
Figure 18: Schematic for the roll, pitch, and yaw movements upon the vehicle that will create a moment.
The fins will produce a rolling moment on the rocket if they are canted at some angle with respect to
the rocket centerline.13 Since the firs for the USA Launch Society’s rocket are not canted and are placed
symmetrically around the rocket body, the rolling moment does not need to be computed since it can be
assumed negligible.11
The pitching moment is the result of the lift and drag forces acting on the rocket. Therefore, the pitching
moment should be considered. The pitching moment can be determined by using the following equation:4
~ θ = Cm qSdjˆ1
M
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(18)
In order to solve for this pitching moment term, the pitching moment coefficient needs to be determined.
This coefficient can be found by using the following equation:11
Cmα =
III.L.6.
2
sinα
[lAb − V olume]
Sd
α
(19)
Wind Drift
For preliminary determinations of wind drift, a very simple mathematical model will be used. In this
model, the parachute is assumed to have no inherent horizontal motion, which in the real world is false, but
the assumption allows for basic analysis of wind drift scenarios. As the recovery section progresses, these
calculations will become more precise and account for more variables.
These calculations will only consider the parachutes drifting after the deployment of the main parachute,
and while it ignores a large amount of potential drift, it allows for more straightforward estimations. The
speed of descent is found using the following equation:10
r
V =
8mg
πρCD D2
(20)
In the above equation, m is 0.45067 slugs, g is gravity, ρ is 0.0023769 slugs/f t3 , CD is 0.8, and D is the
diameter of the parachute which is 6.31758 f t. Inserting these values into the equation, the speed of descent
is determined to be 22.07 f t/s from the point of deployment 800 ft above the ground. This value is further
verified by OPENROCKET determining the same deployment velocity for the simulation. The parachute is
also assumed to have no tendency towards horizontal motion. Using this set of assumptions and determined
values, an estimated value of wind drift can be calculated for various situations.
Under 0 mph wind conditions and the aforementioned assumptions, the rocket would not be expected to
drift at all; although, if it had any horizontal velocity at deployment, it would likely drift in that direction.
Under 5 mph wind conditions, the wind drift could be calculated by dividing 800 by 22.07 to obtain hang
time. Then multiply the hang time by the wind speed converted to f t/s (10 mph = 14.67 f t/s). This
provides a result of 265.82 f t at 5 mph, 531.64 f t at 10 mph, 797.46 f t at 15 mph, and 1063.28 f t at 20
mph.
III.M.
Payload Integration
Two payloads will be integrated with our rocket that we will use in the competition. The first payload will
be a hazard detection camera called a Pixy Cam. An Arduino Uno will be used to power the camera, as well
as log data, after deployment. Both of these components will be attached to a payload tray, firmly nestled
within a payload bay by two metal rods attached to the tray. For extra support needed to safely house
the payload due to the stresses on the rocket body, a firm cardboard coupler will be used as the payload
bay. Each end of the payload bay will be secured by bulkheads made of balsa wood. Balsa wood will be
used for its strength and its light weight. The two metal rods will extend through each bulkhead and be
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secured by hardware. A metal eye-bolt will be attached to the end nearest the nose cone, and will be used to
attach the shock cord. This payload bay will be position below the drogue parachute, located in the booster
bay. A deployment charge will be shot downwards toward the payload bay which will in turn separate the
rocket into two halves. The drogue parachute will then deploy and force the payload bay out of the rocket.
When the payload bay is deployed, the shock cord will be long enough to place the camera well below the
booster bay on descent. Another payload consisting of a direct current turbine will be utilized to harness
power on ascent. This payload will be housed permanently within a separate bay located in the booster
bay above the motor and will be stationary throughout flight. For added stability, the turbine bay will be
positioned between the center of gravity and center of pressure. The same materials will be used for this
turbine bay. The turbine will be held in place within the turbine bay by a special tube-like tray and spokes.
This harnessing apparatus will be made using a 3D printer. Cellulose insulation will be used in between the
turbine bay and motor to protect the turbine components. This bay will be as airtight as possible so no
wind flowing through the bay is lost to the inside of the rocket. On the outside of the rocket body at the
position of the turbine bay, two holes with a half and inch in diameter will be placed above and below the
turbine. The hole above the turbine will use a straight 60 degree angle scoop to direct the airflow toward
the turbine, and the hole below the turbine will be used for the airflow to exit the rocket.
III.N.
Launch Concerns and Operations Procedures
Below are tables that contain the preliminary checklists for each component of the vehicle. Each team
member will be given the opportunity to input any updates to the checklists and verify safety precautions.
As aforementioned, the safety officer will be responsible for the final inspections on each component and his
signature will be provided as well.
Table 11: Checklist for the vehicle’s structural components to ensure excellence before the launch.
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Table 12: Checklist for payloads’ components to ensure security and functionality.
Table 13: Checklist for the assembly of the motor and its components to ensure a safe and successful
launch.
Table 14: Checklist for the recovery system required to have a safe landing and recovery of the vehicle.
Table 15: Checklist for the altimeters to make sure they are all functional especially the scoring altimeter.
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III.N.1.
Final Assembly Checklist
Once all of the preceding checklists are confirmed by the safety officer, the final launch procedures can be
performed.
Table 16: Checklist before the final launch to make sure all preceding procedures are performed.
III.O.
III.O.1.
Safety and Environment (Vehicle and Payload)
Safety Officer
For this year’s competition, the safety officer for the University of South Alabama Launch Society will be
Nghia Huynh. Mr. Huynh will be responsible for the safety of those around him and his job is to ensure that
both the team and the mission are preserved at all times. He currently holds a Tripoli Level 1 certification and
plans on achieving the Tripoli Level 2 certification on February 13, 2015. In order to achieve Tripoli Level 2
certification, the flight must fly a motor in the range of 640.01 and 5120.00 Newton-seconds in total impulse.
Motors within the impulse range are typically J, K, and L motors. There shall be also a written examination
prior to the certification flight in order to evaluate the officer’s rocketry knowledge on the scientific side.
Also, the launch must recover safely in order to be certified.2 With that aside, Mr. Huynh will have to follow
a number of required precautions so that no member or spectator is injured or harmed due to any failures
during the planning, testing, construction, or flights of the rocket. Mr. Huynh will consistently keep the
team informed of any hazards and safety concerns, issues, and documentations throughout the project. He
will create his own personal checklist which will guide the rest of the team to perform the correct pre-flight
and post-flight procedures. Risk matrices will provide the team a way to analyze many of the hazards and
failure possibilities as well as addressing any environmental issues that may be encountered.
Since the safety officer is certified through Tripoli, he will be required to keep all the rocket motors,
explosives, and igniters at all times in a secure, red-painted, plywood-reinforced steel box. The steel box will
be kept off campus due to it being illegal to keep on campus. The only time the box should be unlocked is
when a motor is being removed to be used. Other than to either use or store motors, the box is to remain
closed and secured as all times to avoid any potential explosive hazards.
The University of South Alabama Launch society will obey all rules, regulations, and laws set forth by
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the University of South Alabama, City of Mobile, State of Alabama, Federal Government, and any other
regions that will be encountered during travel. Along with the preceding laws, all traffic and airspace laws
must be obeyed at all times as well. This includes, but not limited to, low-explosives laws and rocketry
associations such as NAR and Tripoli. According to NFPA 1127 code for high power rocketry, a level 2
or 3 flier whose responsibility and duty during the operation of high power rockets is to confirm a rocket’s
compliance with the applicable provisions of this code, be confident that the rocket will fly in a safe manner,
designate the areas of the launch site, and oversee the safety of all spectators and participants.3 Along with
Tripoli regulations, NAR regulations will be obeyed as well. One example for a NAR regulation is the only
use lightweight, non-metal parts for the nose, body, and fins for the rocket.12 Ultimately, the Range Safety
Officer will decide if the rocket is prepared for flight. His decisions and instructions will be respectfully
followed. Along with these regulations and laws, all FAA waiver requirements will be met according to the
time table scheduled.
The highest priority for the University of South Alabama’s Launch Society is the safety of everyone in the
surroundings. The key for this is to be aware of the processes taking place and alert at all times in the event
of hazardous events. Proper personal protective equipment will be worn at all times during the construction
phase of the project. When dealing with any type of chemicals, a fume hood will be used to prevent any
harmful acidic accidents. Black powder, as aforementioned, will be provided by the team’s mentor and will
only be used during test flights. Continuity tests will be performed on electric matches before any flight to
make sure they are working properly in order to ignite the black powder. When dealing with sharp objects
or tools, experienced members will be able to handle the tasks with extra caution. Although everyone is
responsible for their own safety, the safety officer will ultimately be the one to decide if safety guidelines are
being followed correctly.
III.O.2.
Preliminary Hazard Analysis
When dealing with explosives and high power rocketry, risk is inevitably present. The risks encountered in
the project will be approached in the safest way conceivable. In order to prevent hazards, risk matrices will
be provided and used to display the level of danger for different scenarios. Creating a risk matrix will allow
the team to update the information to the chart as different situations are being encountered; therefore,
preventing the same mistakes from reoccurring.
III.O.3.
Personnel Hazards
It is essential to recall that in order for the project to be a success, both the team and rocket must be
successful. The team members must stay safe and unharmed while the rocket must conquer the competition
guidelines for success. For the purpose of the following risk matrices, both the team and rocket will be
considered as the personnel for the project. The risk matrices that pertain to the personnel of this project
are shown in Tables 17 and 18.
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Table 17: Risk matrix for the potential hazards that can occur during travel.
• Deadlines (High)
Deadlines for this project are designed to accommodate proper planning and time management. These
deadlines must be promptly met and it ensures that no agreements are broken, ranking it high in
importance. In order to accomplish these deadlines, time will be set aside for each task in order to
provide a cushion for the team to fall back on if needed. This will improve productivity and compensate
for any error.
• Prolonged sun exposure (High)
Whether it is during test flights or the launch competition, sun exposure is inevitable to all participants.
In the past, the radiation caused great dermal irritations and potentially skin damage due to the long
hours of exposure. It is extremely important that the team stay protected by wearing sunscreen and
properly applying it constantly.
• Traffic laws (Medium)
Different or modified traffic laws are expected to be encountered during the traveling to test and fly
the team’s rocket. Due to these rules and regulations, the team must drive respectfully and be aware
at all times of the constant changes, especially in unfamiliar areas. If this hazard is handled correctly,
the hazard would be considered a low risk.
• Travel fatigue (Medium)
During the stretch of the project, the team is expected to be able to travel anywhere necessary in
order to test launch the rockets. This will imminently lead the team exposed to fatigue and maybe
drowsiness if experienced for extended time periods. It should be noted as a concern for members of
the team as the hazard of being fatigued could lead to accidental events.
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• Flat tires (Low/Medium)
During road travel, flat tires have a possibility of occurring. It will be a minor hazard if the team
is aware and properly prepared. Making sure that spare tires are available and inflated will prevent
problems from becoming major. All tires should be inspected and confirmed to be in the best condition
possible.
• Time delays (Medium)
Events may occurs or arise that may possibly cause time delays for the team. It has the chance of being
a moderate hazard as it could create other problems later on. If the team follows orders, checklists,
deadlines, and stays ahead of their assigned tasks, the chance of time delays causing an issue should
be minimal. To assist in preventing time delays, likewise with deadlines, extra time will be allotted to
events and travel to compensate for errors.
• Broken or lost rocket components (High)
As the team is traveling from various locations, packaging and handling issues may be present. In
order to mitigate losing or breaking any items, extra provisions will be taken. Every component of the
rocket must be labeled and packaged for protection. Sensitive components should be properly padded
while components such as the body tubes are held stationary in a rack or case. The created checklist
in the preceding safety section will be used in order to prevent this hazard.
• Loss of transportation (High)
Since the team is driving and transporting the rocket to Huntsville, Alabama, there is an improbable
chance that the vehicle will break down. Therefore; this can be devastating and cause a large delay
and negative impact towards the team budget. To mitigate this hazard, proper engine maintenance
will be administered such as checking oil, fluids, gauges, etc. Other arrangements that can be possibly
affect the budget will be considered during budget planning. This calls for the team to be alert of the
transportation vehicle’s condition at all times.
• Separation from the group (Medium)
The loss of a team member is a rare occurrence that nevertheless is considered a hazard to the project.
Prevention can be easily implemented through the use of cellular phones to keep in contact with one
another via calling or text messaging. Also, team members must be accompanied by at least one other
member to ensure safety.
• Dust inhalation (Medium)
Dust inhalation will be encountered when sanding body tubes of the team members certification rockets.
Wearing dust masks or vacuum hoods to reduce the chances of breathing the material will make this
a medium concern, yet it could cause minor respiratory issues if not followed.
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Table 18: Risk matrix for the potential hazards that can occur during the construction phase.
• Minor abrasions from sandpaper (High)
It should noted that the team will use sandpaper and teammates should wear gloves when utilizing
this tool. This can cause damage to members’ hands and result in delay in production.
• Use of machinery and tools (Medium)
The majority of the construction that concerns the use of machinery and tool will be performed by
a member that has had experience. This is to enforce the mitigation of any damage to the items,
components, or people. Otherwise, proper lessons will be administered to the inexperienced.
• Sharp knives (High)
The use of sharp knives and objects are to be expected. Team members will handle these tools maturely
to prevent any unnecessary harm. Team members will use these items while wearing protective gear
such as gloves, protective eye wear, and proper precautions.
• Contact with paint (Low/Medium)
When dealing with paint, members must be considerate of the surrounding and the environment they
are painting in. Gloves, proper clothing, and protective eye wear will be worn to prevent any contact
with members’ skins.
• Possibility of contact with glue (Medium)
Glues will be used to seal components of the rocket. Wearing gloves should prevent contact. However,
if it does occur, immediate cleaning will be performed to prevent any subsequent consequences.
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• Chemical exposure (High)
At any time a chemical is being used by a team member, MSDS’s will be nearby in all laboratories
and online and will be reviewed to follow proper precautions for safe use and how to apply protective
equipment. This could severely harm an individual if not carefully executed. It is possible that team
members will at some point handle a chemical and should do so carefully and with a partner to reduce
risk.
• Contact with fiberglass and carbon fiber resin (Low/Medium)
A couple of the team members may test with fiberglass and carbon fiber. This may require the use
of resins to set the fibers. Wearing gloves, proper clothing and eye-wear will diminish the chance of
contact.
• Accidental discharge of black powder (Medium)
Only certified members will be allowed to contact black powder, making it improbable that any member
would encounter it. However, team members may be near when black powder is being utilized. This
presents an chance of an accidental firing of black powder around people that could be catastrophic
under the right conditions. Keeping team members at safe distances from all charges and staying aware
will reduce the chance of harm and failure.
III.O.4.
Failure Modes
The following tables below display the risk matrices for the possibility of failures present in the project. The
failure modes are expressed in color to display the magnitude of each possibility. The failure modes that will
be covered are for the proposed design of the rocket, payload integration, and launch operations. Also, an
explanation for the mitigation of the hazards will be provided as well.
• Minor damage to shock cord from ejection charge (Medium)
Any damages to the shock cord can be easily prevented by selecting the proper shock cord material
that will be highly resistant to high temperatures. As the ejection charge activates, corrosive gases will
emit and can potentially damage the shock cord; therefore, the success of the project is diminished.
• One altimeter fails (Low)
This possibility of failure can be avoided by simply having extra altimeters ready to replace any that
fails. However, a more efficient way to prevent this failure from happening is to properly secure the
altimeters onto the vehicle wherever the altimeter may be.
• Minor damage to parachute from ejection charge (Low/Medium)
The parachute can experience possible damage upon deployment stages. Proper insulation will be
applied to protect the parachutes on the vehicle in order to recover the vehicle safely.
• Premature ignition of ejection charges (Medium)
During the flight, abrupt changes in pressure could possibly cause altimeters to receive false readings
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Table 19: Risk matrix for any hypothetical failures that can be incurred upon the vehicle.
and deploy at the incorrect moment. This would potentially cause sever damage to components. By
creating a streamline rocket exterior, the pressure should remain consistent around the port holes. The
selection of the quality of the altimeters will also aid in the mitigation of this failure.
• Major damage to parachute from ejection charges (High)
If improper packing is done on the parachutes or protective material, major damage could arise from
the incident. Therefore, this can result in the parachute failing to deploy properly and the vehicle will
not have a safe landing. To prevent this catastrophic failure, the parachute packing will be shadowed
by our team mentor and safety officer to ensure a successful descent.
• Scoring altimeter fails (Medium)
By limiting the function of the scoring altimeter to just recording data, the team hopes to reduce the
chance of it failing to improbable. It will have a dedicated power source and tested for accuracy. If it
fails, replacements will be be available and additional launches will be performed.
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• No separation from ejection charge (High)
If there is no separation from ejection charge, the launch will be considered ballistic and a failure.
This could harm both spectators and team members because the behavior of the launch would be
arbitrary. By testing for adequate amounts of black powder, the team expects to determine the proper
amount needed to create the desired separation of rocket segments. If the launch fails to replicate test
results, the mission disaster would be catastrophic. With enough testing, this failure mode should be
improbable.
• Shear pins do not fail properly (High)
Proper testings will be utilized to establish security to the limited number of shear pins to restrict the
rocket from falling apart. Also, it can fall apart separately if the team desires. Failure to break the
shear pins will result in the rocket being eradicated.
• Fins separate from rocket (Medium)
The fins of the rocket are the primary part to create a turbulence on the rocket. The vehicle will
experience violent behavior which could separate the fins from the rocket. Reinforcing the fins will
add rigidity to the structure and allow a proper launch. One way of reinforcing the fins is applying the
proper adhesive and allowing days for it to fully be applied.
• Failure of both altimeters (Medium)
In the event the deployment altimeters fail to set off the ejection charge, the rocket would become
ballistic and catastrophic. No deployment would occur and the rocket could possibly be destroyed.
Having the altimeters with their own dedicated power source and electric match is designed to reduce
this failure.
• Corrosive gases from ejection charges (Medium)
As the ejection charges are activated, harmful gases are released as an aftermath effect and can consequently damage electronics, wires, batteries, and other sensitive components. Having secured and
sealed bulkheads between each compartment is a precaution the team will undertake to tolerate failure.
• Explosion from ejection charge (High)
By performing tests with the black powder charges, the goal is to determine the correct amount to
separate the rocket segments without destroying the rocket. Calculations have been done as a starting
point to find an amount to begin with; however, further testings are necessary in order to verify.
• Corrosive gases from motor (High)
Damage from the motor is possible with every launch. Properly fixing the motor assembly should allow
the motor to expel all gases from the aft end. Creating another bulkhead will also help to seal the
compartment.
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Table 20: Risk matrix for any hypothetical failures that can be incurred towards the payload.
• Power to equipment is lost (Low)
If the power is lost to the equipment, the payload for the mission will not be able to detect any ground
threats. However, this possibility will not affect anyone’s safety; rather, it will only affect the success of
the launch team since it is a required payload. Issuing redundancy in the form of extra power supplies
will create a comfort zone for this failure mode.
• Explosion from motor (High)
The team has concurred that a manually packed motor will be used since a reusable motor tube is
available for use. Choosing this method will open the gates for sources of failure occur such as an
incorrectly packed motor that maybe is missing a retention ring. Though improbable, this explosion
would be catastrophic and potentially destroy the payload. Choosing a reliable supplier and having
the team mentor examine the motor will eliminate this potential failure.
• Activate altimeter (High)
In the event that the rocket is bumped or knocked over once the altimeters are on, the ejection charges
could ignite. This has the potential to be extremely dangerous to others in the area. Utilizing checklists
and prioritizing steps in the launch process aims to reduce this risk. The extended exposure to sunlight
could also result in the activation of altimeters. By having a non-transparent case, the team aims to
prevent sunlight from reaching the inside of the bay.
• No liftoff (Low)
In the rare occasion where a launch system is incorrectly connected, a no launch situation may occur
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Table 21: Risk matrix for any hypothetical failures that can be incurred during the launch.
and the rocket motor does not activate. This is a minor mode of failure that should not endanger
personnel. It can be easily resolved by rechecking the electric matches and motor components.
• Faulty motor (Medium)
Having selected a reputable motor, the chance of the motor failing is diminished. However, in the rare
instance that the motor explodes, the result would be detrimental and failure will be present.
III.O.5.
Environmental Hazards
While technical hazards is the more common form of hazards, environmental hazards are present as well.
As the project progresses, the environment and weather will become a developing issue. For example,
appropriate weather is required in order for test launches to launch due to airspace regulations. These
events could compromise part or all of the mission, depending on the magnitude of the situation and how it
unfolds. Therefore, time management and efficient planning is required in order to avoid any environmental
hazards. Keeping up to date with current events and news will also aid to reduce this hazard. Provided
below is the risk matrix for any potential environmental hazards that can be present to the team.
• Prolonged sun exposure (High)
The team will encounter periods of extended exposure to the sun while in the open fields of Huntsville.
During the extended periods of sun exposure, ample amounts of sunscreen will be applied and proper
clothing will be worn to mitigate this issue. If appropriate precautions are executed, this hazard will
be of no concern. Not only will the radiation be an issue to team members, it will also be an issue for
the rocket. Many altimeters contain barometric sensors which may be damaged due to prolonged sun
exposure. By implementing an altimeter bay that is opaque, the team will be able to prevent the sun
from damaging any altimeter and electronic components.
• Heat (Medium)
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Table 22: Risk matrix for any hypothetical hazards and concerns that the environment may bring.
Due to the time of the year of the launch, heat will need to be addressed. Members should try to
remain cool and avoid excessive heat exposure, which could lead to heat strokes. Taking breaks from
the heat will help keep the team safe. Also, team members should watch other for signs of a heat
stroke such as flushed skin, rapid breathing, and headaches. The heat may also create concerns for
equipment in the rocket. For this reason, sensitive parts should be kept in a cool environment until
needed.
• Dehydration (High)
As the sun and heat impact the team, it is important to remain hydrated at all times. There will also
be refreshments available at all times to relieve of this risk.
• Cold (Low/Medium)
The climate for Huntsville during the time of the competition will pose little to no threat for cold
weather conditions. Although a minor risk, team members will be required to pack at least some
clothing to resist the cold weather. As the rocket travels upward in the atmosphere, it will be reaching
cooler environments. Precautions have been made to verify the compatibility of rocket components
and what temperature it can operate under.
• Rain (Medium)
The chance of rain may come at any time during the period of this project. Team members are expected to have either rain jackets or umbrellas to remain dry. Along with the members, the equipment,
especially electronics, must remain dry as to not cause any deformation of any rocket segments or damage to any sensitive electronics. The rain may also create delays which may need to be accommodated
for as they are encountered. Weather forecasts will be closely observed during any events in which the
team may partake.
• Moisture (High)
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Properly sealing all electronics and equipment will be essential for the success of the mission. Corrosion
or deformation of parts due to moisture could cause devastating results to the launches if not prepared
for. Parts will need to be coated in water resistant material to reduce the chances of failures. All
explosives will also need to be properly maintained in dry storage at all times. Any electronics used
will be contained in a dry environment until used.
IV.
Payload Criteria
Table 23: Payload purpose and description
Payload
Pixy Camera (CMUcam5)
Payload Requirement
X, Y, and Z coordinates
for hazard detection
Subsystems
Arduino Uno, brightly colored tarps, 9V battery,
Stratologger altimeter
DC Motor Turbine
Create at least 2V during
half the time to apogee
Arduino Uno and 9V Battery
IV..1.
Analysis
Identify the exact location
by association with the
height given by the altimeter
Provide enough charge
from the turbine to theoretically power an e-match
at apogee
PixyCam
Table 24: Camera selection criteria.
Camera Module
Pixy Camera (CMUcam5)
Weight
0.1 lb
Dimensions
2.2 x 1.5 x 2.2 in
Cost
$69.00
Application
Object Detection and Tracking
Using the PixyCam and an Arduino Uno and its respective IDE, the relative x and y coordinate positions
of certain objects can be tracked and logged. Before launch, we will use the Pixy Cam software to capture
a photo of at least three distinguishable colors. These colors will have to differ from the surrounding launch
environment to ensure optimal detection. The software will verify these colors and set them as unique IDs.
As data is logged to the Arduino Uno, the relative size and positions can be tracked through descent, as this
payload will be deployed at apogee. Large tarps with the unique colors discussed will be placed in random
positions around the launch site for detection. Upon completion of the flight of the rocket, the logged data
will be uploaded to a computer for analysis. We will then integrate the coordinates of the objects from the
Pixy Cam with the height of the rocket at that time from the Stratologger altimeter. Using this information,
exact position with minimal tolerance will be extracted. The x-coordinate of the object can be obtained by
Equation (21).
x = ztan(γx )
(21)
Where x is the actual x-coordinate of the detected object, z is the height of the rocket, and γx is the
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x-direction angle of actual view to be calculated. The y-coordinate of the detected object can be obtained
with Equation (22).
y = ztan(γy )
(22)
Where y is the actual y-coordinate of the object and γy is the y-direction angle of actual view to be
calculated. γx can be determined using Equation (23).
γx = arcsin
xγ sin(α)
xα
(23)
Where xγ is the x camera coordinate given by the Pixy Cam, α is the field of view angle tested (37.5◦ ),
and xα is the maximum x-coordinate of the Pixy Cam. γy can be found using Equation (24).
γy = arcsin
yγ sin(β)
yβ
(24)
Where yγ is the y camera coordinate given by the Pixy Cam, β is the field of view angle tested (37.5◦ ),
and xβ is the maximum y-coordinate of the Pixy Cam.
IV..2.
DC Turbine
Figure 19: Example wind turbine that will be used as a payload.
As for the second payload, a dc motor will be used as a turbine to provide power to a hypothetical drogue
deployment source. A propeller will be permanently attached to the dc motors shaft. Air will be directed
into the upper portion of the turbine bay through a hole extending through the rocket body to the interior
of the turbine bay. In order to gain as much air as possible through the provided hole, a small scoop that
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is pointed upwards will be secured to the rocket body just below the outside top hole as shown in Figure
21. Air gathered into the turbine bay will escape through a hole located at the bottom of the turbine bay
just below the turbine. These holes will extend through the turbine bay to the outer portion of the rocket
body. The dc motor will be firmly attached to the turbine bay with a turbine tray as seen in Figure 20.
An Arduino Uno will be used with a data logging shield to capture the voltage provided by the turbine on
ascent. The ideal voltage to be obtained is provided in the dc turbine theory section.
Figure 20: Isometric view of the turbine tray used to hold the turbine into the rocket body.
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Figure 21: Air scoops used to provide air to the turbine.
IV.A.
IV.A.1.
Payload Testing
PixyCam
The accuracy of the position will depend on the relative positions on the camera. As the camera moves
closer towards the tarp, the tarp will appear larger on the screen. The PixyCam shows the position of the
center of the object; therefore, if the tarp is close enough to the camera, then the camera will show a position
of (0,0) because the tarp will take up the whole screen of the camera. Testing and accuracy of the camera
will be accomplished by creating a unique color object and measuring the distance from the object to the
camera. We will then use the above equation to test whether the data received from the Pixy Camera and
Arduino are accurate and correct. Refer to the Test 1 and Test 2 results below for the calculated results.
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Table 25: Test results for PixyCam using the field-of-view equation.
xactual (in)
yactual (in)
zactual (in)
xγ
yβ
xα
yβ
α
β
xcalculated (in)
ycalculated (in)
IV.A.2.
Test 1
12
29
228
14
20
150
100
37.5◦
37.5◦
12.975
18
Test 2
1
16
34
6
80
150
100
37.5◦
37.5◦
0.8217
11.671
DC Turbine
Testing will be performed prior to launch to provide feasibility for the payload. First we will test the turbine
ft
voltage output in a wind tunnel set to the highest speed of launch ascent (678 sec
) given in OpenRocket.
Next we will put together the turbine bay along with the scoops and holes of the specified size. The turbine
will be secured in the bay by the turbine tray. We will then test the voltage output of the system in the
wind tunnel given the highest velocity of ascent given in OpenRocket.
We have not received the final dc turbine that we will use in the final competition. However, we have
done testing with a scaled down version of the final turbine. We produced a voltage of 1.15 V with this
turbine from an air nozzle in South Alabama’s Aerospace Propulsion Lab.
IV.B.
IV.B.1.
Uniqueness and Difficulty of Payloads
PixyCam
Each of these payloads provides a uniqueness to the rocket that will provide added safety and improved
recovery. When a launch vehicle descends back to the ground, it may be holding valuable material or data.
Recovery of the launch vehicle is of the utmost importance to obtain the data that is logged on the rocket
for analysis or the obtained luggage and to assess the launch vehicles condition. The recovery system relies
heavily on the hazard detection system of the launch vehicle. If the vehicle were to land out of reach,
such as on a mountain, recovery becomes difficult or impossible. Electronics on board the vehicle may be
damaged by certain environments such as extreme heat or water, and should be guided away from these
hazards to recover the data on board. The Pixy Camera will provide valuable feedback of the relative x and
y coordinates of these hypothetical hazards by way of color recognition. An exact position of these hazards
can be found by correlating its x and y data with the height (z) data from the altimeter. After flight, the
data will be extracted from the Arduino Uno using a SD card shield that logs the data during flight. The x
and y positions at different points in time from the Pixy Camera will then be loaded into Excel along with
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the height at different points in time from the altimeter. By using this equation, we will be able to find the
exact positions of the hazards relative to the camera. This payload will provide a small challenge by using
the data and equation to find the position of the hazards.
IV.B.2.
DC Turbine
At apogee, a drogue parachute must be deployed to slow the launch vehicle to a speed that is suitable for a
main parachute deployment. Safety of the rocket, surrounding objects, and surrounding people depend on
this drogue deployment. Usually these drogue deployments are powered by an on-board altimeter and 9v
battery. If the battery or altimeter were to malfunction, a serious safety hazard will have been formed from
the descending vehicle. The vehicle will accelerate and the main deployment may zipper the vehicle. If the
main deployment fails due to this zippering, the vehicle will plummet to the ground at astounding speeds.
Providing a turbine power source that is harnessed within the rocket body will prove to be a difficult task
given the chaotic fluid dynamics of the system.
IV.C.
Turbine Theory
As power is generated by the turbine, the voltage will be harnessed within a capacitor. A resistor in series
with the capacitor will provide a simple RC circuit for storing power. The RC circuit will charge in accordance
with Equation (25).
−t
Vc = Vs (1 − e RC )
(25)
Where Vc is the voltage across the capacitor, Vs is the voltage from the source, t is time, R is resistance,
and C is capacitance. A switch, possibly a p-type MOSFET transistor, will be connected in series directly
after the capacitor. This transistor or switch will only allow current to flow when the voltage from the
turbine reaches zero. To avoid current leaking from the capacitor back into the motor, a current directing
diode will be attached in series with the resistor. Another resistor of low resistance (nichrome wire) will be
attached in series with the switch to provide a current large enough to power and e-match. The voltage from
the capacitor will be discharged according to Equation (26). As the voltage discharges, it will pass through
the resistor providing a current.
−t
Vc = Vs e RC
(26)
The e-match consists of a black powder charge. Black powder charge has a flash of approximately 801867◦ F. In order to obtain a temperature of 1000◦ F needed to ignite the black powder charge, a current of
1.78 amps through nichrome wire is needed. Providing that the voltage supplied by the dc turbine is at least
2 volts, the resistance needed to obtain this current can be found through Ohms Law, shown in Equation
(27).
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I=
V
R
The resistance found is 1.1236 Ω. Resistance of 30 gage copper wire was found to be 0.339
(27)
Ω
m;
thus with
4 inches of copper wire used the resistance is 0.034442 Ω. Resistance of nichrome wire was found to be 6.75
Ω
ft .
Nichrome used with a length of 1 inch would be 0.5625 Ω, providing a total resistance of 0.59694 Ω.
Using Ohm’s law, the current given with the voltage of 2 volts and resistance of 0.59694 Ω would be 3.3504
amps, which would be ample current to ignite the black powder charge.
Ideally, the capacitor should be 66% charged in half of the time it takes the rocket to reach apogee (8.5
seconds). The size of the capacitor can be calculated with Equation (28).
c=
−ta
Rtot ln(0.34)
(28)
Where c is capacitance, ta is half the time to apogee (8.5 s), and Rtot is the total resistance (.59694 Ω).
The capacitance found was 14.007 F.
Figure 22: Proposed circuit schematic for the e-match
The second payload of a turbine could theoretically provide a backup system for the drogue deployment.
Based on the circuitry seen above, the turbine will be activated by the outside wind on ascent. This voltage
created by the turbine will charge a capacitor. A switch will prevent the capacitor from discharging. The
switch will be able to detect if the turbine is creating a voltage. If a voltage is being created, the switch will
remain open. If there is no voltage being provided by the turbine, the switch will close and the capacitor
will then discharge. Our hypothesis is that the turbine will cease to spin at apogee when there is little to
no airflow through the chamber. At this time the capacitor will discharge the voltage and create a current
from the resistor.
IV.D.
IV.D.1.
Payload Concept Features and Definition
Creativity and Originality
Each of these payloads provides a uniqueness to the rocket that will provide added safety and improved
recovery. When a launch vehicle descends back to the ground, it may be holding valuable material or data.
Recovery of the launch vehicle is of the utmost importance to obtain the data that is logged on the rocket
for analysis or the obtained luggage and to assess the launch vehicles condition. The recovery system relies
heavily on the hazard detection system of the launch vehicle. If the vehicle were to land out of reach,
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such as on a mountain, recovery becomes difficult or impossible. Electronics on board the vehicle may be
damaged by certain environments such as extreme heat or water, and should be guided away from these
hazards to recover the data on board. The Pixy Camera will provide valuable feedback of the relative x and
y coordinates of these hypothetical hazards by way of color recognition. An exact position of these hazards
can be found by correlating its x and y data with the height (z) data from the altimeter. After flight, the
data will be extracted from the Arduino Uno using a SD card shield that logs the data during flight. The x
and y positions at different points in time from the Pixy Camera will then be loaded into Excel along with
the height at different points in time from the altimeter. By using this equation, we will be able to find the
exact positions of the hazards relative to the camera. This payload will provide a small challenge by using
the data and equation to find the position of the hazards. The accuracy of the position will depend on the
relative positions on the camera. As the camera moves closer towards the tarp, the tarp will appear larger
on the screen. The Pixy Camera shows the position of the center of the object;therefore, if the tarp is close
enough to the camera, then the camera will show a position of (0,0) because the tarp will take up the whole
screen of the camera. Testing and accuracy of the camera will be accomplished by creating a unique color
object and measuring the distance from the object to the camera. We will then use the above equation to
test whether the data received from the Pixy Camera and Arduino are accurate and correct. After the data
is received, we will then be able to provide a tolerance.
At apogee, a drogue parachute must be deployed to slow the launch vehicle to a speed that is suitable
for a main parachute deployment. Safety of the rocket, surrounding objects, and surrounding people depend
on this drogue deployment. Usually these drogue deployments are powered by an on-board altimeter and 9v
battery. If the battery or altimeter were to malfunction, a serious safety hazard will have been formed from
the descending vehicle. The vehicle will accelerate and the main deployment may zipper the vehicle. If the
main deployment fails due to this zippering, the vehicle will plummet to the ground at astounding speeds.
The second payload of a turbine could theoretically provide a backup system for the drogue deployment.
Based on the circuitry seen above, the turbine will be activated by the outside wind on ascent. This voltage
created by the turbine will charge a capacitor. A switch will prevent the capacitor from discharging. The
switch will be able to detect if the turbine is creating a voltage. If a voltage is being created, the switch will
remain open. If there is no voltage being provided by the turbine, the switch will close and the capacitor
will then discharge. Our hypothesis is that the turbine will cease to spin at apogee when there is little to
no airflow through the chamber. At this time the capacitor will discharge the voltage and create a current
from the resistor. This current will have to be large enough to theoretically power an e-match. Testing will
be performed prior to launch to provide feasibility for the payload. First we will test the current output of
the Stratologger altimeter. Next we can test the turbine voltage output in a wind tunnel set to the speed of
launch ascent given in OpenRocket. If the voltage is high enough, the turbine will be attached to the circuit
and run in the wind tunnel. We will stop the wind from the wind tunnel after the time to apogee given
in OpenRocket and measure the current provided by the capacitor discharge and resistor. If this current is
close to the current measured by the altimeter, the turbine will be a feasible payload.
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IV.D.2.
Suitable Level of Challenge
Both payloads will present its own challenges and difficulty. As for the Pixy camera, maintaining a reasonable
precision and accuracy will be rather difficult with elevation changes at ground level, barrel distortion effects
of the camera, camera tilt, and the camera swing. The turbine will have to generate enough voltage to heat
a small resistive wire hot enough to light an E-match. A reasonable voltage will have to be obtained within
the time it takes the rocket to launch to apogee. These unique challenges and difficulties will only make the
team strive harder to reach set goals and obtain success.
IV.E.
Testing and Design of Payload Equipment
IV.F.
Science Value
V.
V.A.
V.A.1.
Project Plan
Budget Plan
Rocket Structure Budget
In the following tables, all the essential materials and components relating to the vehicle are listed along
with reasoning and prices.
Table 26: Rocket structure budgeting.
Rocket Structure
Airframe
Price
$120
Fins
Motor Set
$28
$113
Center Rings
Miscellaneous
$20
$100
Subtotal
$381
Reasoning
The team is budgeting $120 for the airframe including
the nose cone.
$28 will be budgeted for the construction of the fins.
Complete motor - includes one casing, one nozzle, one
forward bulkhead, two snap rings, and one nozzle washer.
$20 will be used for the center rings.
The team is adding $100 for miscellaneous items possibly
needed for the airframe.
$381 is the total cost of the airframe.This budget is
currently tentative with respect to the material down
selection which will be presented in the PDR.
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Table 27: Recovery components budgeting.
Recovery Components
E-matches
Cable Cutters
Main Parachute
Price
$10
$55
$225
Drogue Parachute
$6
Black Powder
Shock Cord
$20
$10
Miscellaneous
Sewing Machine
Subtotal
$100
$0
$426
Reasoning
There will be a set $10 budget for e-matches.
Used for dual deployment purposes from the same bay.
The team will manufacture the main parachute with
the given budget.
The team will manufacture the drogue parachute with
the given budget.
There will be a budget of $20 for black powder.
Minimal cost due to the team having an ample amount of
shock cords from previous launches.
Flexible amount in the event of emergency.
Members of the team possess a sewing machine.
The projected budget strictly for recovery components.
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Table 28: Altimeter bay budgeting.
Altimeter Bay
Altimeter(x2)
Shear Pins(x2)
Price
$140
$8
Misc. Components
$100
E-match Material
Bay Casing
$100
$70
Ejection Charge
$130
Subtotal
$548
Reasoning
The team will be using two Perfectflite Stratologgers altimeter for the vehicle.
Both 2-56 and 4-40 shear pins are available in the Aerospace
Propulsion Lab.
Extra parts and supplies such as switches, screws, bolts,
batteries, glue, etc.
Includes prefabricated wires or wires and ignition.
The team may build the altimeter bay from scratch. However,
LOC/precision is being considered being that it is available
in different diameter tubing.
Ejection charging material needed to separate the sections
of the rocket includes black powder, pyrodex, and CO2 kits.
The subtotal describes the range of price for the components
and parts for the altimeter bay to be between $548.
Table 29: Certification budgeting.
V.A.2.
Certifications
Certification Kits (x2)
Price
$200
Certification Motors (x2)
Subtotal
$160
$360
Reasoning
The team will budget $200 for certification kits
to be flown September.
The team is budgeting $160 for certification motors.
The total cost of the certification is $360.
Outreach Budget
The team recognizes the involvement of young students as a fundamental part of the project. For this
reason, the team is working together with outreach specialist Kelly Jackson to participate in events that
can motivate students to study mathematics and science. Therefore, the team has prepared a budget to
participate in this activities. The budget is illustrated below.
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Table 30: Outreach budgeting.
V.A.3.
Outreach
Model Parachutes (x10)
Price
$75
Straws
$10
Gas per Visit (x7)
$100
Rocket Motors (x10)
$200
Subtotal
$385
Reasoning
The team is budgeting $75 that is divided into 4 yards
of nylon and lines for the parachutes.
$10 is budgeted to buy regular straws for a
pressure rocket system.
$100 is designated for gas to attend the outreach
activities.
$200 is budgeted for the mini rockets that will
be launched at schools.
The subtotal cost for the the outreach activities is $385.
Budget for Trips to Samson, Alabama
The budget for the trips to Samson, AL is fundamental for both testing and evaluation of the rocket. The
trips are going to allow the team to launch the rockets with the payload and perform the test for the entire
time of the project. The budget overview relative to this is provided below:
Table 31: Budget for traveling to Samson, AL.
V.A.4.
Travel to Samson, AL
Samson, AL (Driving x10)
Price
$900
Subtotal
$900
Reasoning
The team is projected to attend a potential of 10
available launches. One will be held once every 3
weeks starting September 12, 2015.
Total for transportation to launches in Samson, AL.
Huntsville Budget Plan
In the following table, the team presents a budget plan to travel to Huntsville, AL. The team will drive
approximately 5 hours to Huntsville, AL and stopping one time per trip. Hotel expenses are also budgeted
for staying in Huntsville, AL for 3 to 4 days.
Table 32: Budget for trip to Huntsville, AL.
Travel to Huntsville, AL
Driving (3 cars)
Price
$405
Food (x5)
Hotel (3 nights)
$300
$700
Subtotal
$1405
Reasoning
In order to budget for the trip to Huntsville, AL, there
will be 3 vehicles taken. It will be an estimate of $90
per car for round trip, and an additional $45 for travel
within the city.
$300 is budgeted for the teams food expenses.
The team is budgeting $700 for hotel expenses for
approximately 8 people.
The total for travel cost to Huntsville, AL considering
all the gas, food, and lodging prices.
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V.A.5.
Total Project Cost
Table 33: Total project budgeting.
Total Project Cost
Rocket Structure
Payload
Recovery Components
Altimeter Bay
Certifications
Outreach
Travel to Samson, AL
Travel to Huntsville, AL
Total Project Cost
V.B.
Price
$381
$400
$426
$548
$360
$385
$900
$1405
$4806
Funding Plan
The funding plan is a fundamental aspect for the project to be a success. Without any funding, it is nearly
impossible for the team to have a budget for the vehicle to be fully constructed. For this reason, the team has
requested funding from three major organizations: the Student Government Association (SGA), Alabama
Space Grant Consortium (ASGC), and other local companies such as Airbus. The Student Government
Association will be able to provide $2000 for the project plus an additional $1000 for travel expenses. In
addition, the team has obtained a grant from the Alabama Space Grant Consortium (outreach program) for
$5000 to provide for conferences, seminars, and supplies that can contribute with the promotion of science
for young students in the area. The team has also contacted local companies for sponsorships that can help
the project become a success.
V.C.
Project Timeline
This section will cover the vital events the team will encounter that involves with NASA deadlines and
launch days. A Gantt chart will also be provided in order to give a visual organization of team events.
V.C.1.
Event and Submission Schedule
Table 34: Event and submission schedule.
August 2015
Event
7
Request for Proposal (RFP) goes out to all teams.
September 2015
11
Electronic copy of completed proposal due to project office
by 5 P.M. Central Time.
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11
Finalize Inventory for SEARS Launch.
12
SEARS Launch in Samson, AL.
October 2015
2
Awarded proposals announced.
3
SEARS Launch in Samson, AL.
7
Kickoff and PDR Q&A.
23
Team web presence established.
November 2015
6
Preliminary Design Review (PDR) reports, presentation slides,
and flysheet posted on the team website by 8:00 A.M. Central Time.
7
SEARS Launch in Samson, AL.
9-20
PDR video teleconferences.
December 2015
4
Critical Design Review (CDR) Q&A.
5
SEARS Launch in Samson, AL.
January 2016
15
CDR reports, presentation slides, and flysheet posted on the
team website by 8:00 A.M. Central Time.
16
SEARS Launch in Samson, AL.
19-29
CDR video teleconferences.
February 2016
3
Flight Readiness Review (FRR) Q&A.
6
SEARS Launch in Samson, AL.
27
SEARS Launch in Samson, AL.
March 2016
14
FRR reports, presentation slides, and flysheet posted to team
website by 8:00 A.M. Central Time.
19
SEARS Launch in Samson, AL.
17-30
FRR video teleconferencses.
April 2016
5-7
AIAA Student Conference in Starkville, MS.
13
Teams travel to Huntsville, AL.
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13
Launch Readiness Reviews (LRR).
14
LRRs and safety briefing.
15
Rocket Fair and Tours of MSFC.
16
Launch Day.
17
Backup Launch Day.
29
Post-Launch Assessment Review (PLAR) posted on the team
website by 8:00 A.M. Central Time.
May 2016
11
Winning team announced.
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V.C.2.
Gantt Chart
Figure 23: Gantt chart showing the projected schedule for the team to complete competition tasks.
V.D.
V.D.1.
Educational Outreach
Purpose
As actively involved members of the local community, the society’s goal is to share our knowledge, interests,
and experience to the collective body of the Gulf Coast and South regional area. With the expansion of
intellectual knowledge as our motivation, we week to engage middle school and high school students in
learning to promote higher education in the fields of mathematics and science. Elementary students will
have an opportunity to expand their interest in model rocketry and science that will gently assist them to
being exposed to simplified engineering principles. To correspond with the USLI competition, USA Launch
Society’s outreach events will focus on the fundamental concepts of rockets, scientific principles of rocketry,
and allowing the pupil to get hands on experience with model rockets.
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V.D.2.
Status of Outreach
Currently, the team has attended two total outreach events to two different schools in Brewton, Alabama.
From the two schools, T.R. Miller High School and W.S. Neal Elementary, a total of 204 students have been
outreached to. The team will be proactive in attending and preparing for the future outreach events that will
be listed in 35. The team will be collaborating with the team’s high school intern, Kayla Bell, throughout
the remainder of the semester in order to plan an additional outreach event at the Alabama School of Math
and Science. For the outreach events, the team will be purchasing model rocket kits from a local hobby store,
HobbyTown USA, that will be able to sustain up to level D rocket motors. The team will be demonstrating
the different levels of impulse and the significant difference in the ascending levels. Figures 24 and 25 are
pictures captured during the outreach that the team performed.
Figure 24: Andrew Tindell (left), Nghia Huynh (middle), and Conner Denton (right) teaching the students
at W.S. Neal Elementary the fundamentals of rocketry.
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Figure 25: Conner Denton (left) preparing igniters for Andrew Tindell (right) as Andrew is setting up the
launch rod for the kit rocket.
V.E.
Schedule for Outreach Events
Table 35: Outreach events schedule.
Date
November 13, 2015
December 14, 2015
January 19, 2016
February 11, 2016
February 15-19, 2016
February 26, 2015
Location of Outreach
T.R. Miller High School in Brewton, AL
W.S. Neal Elementary School in Brewton, AL
Alabama School of Math and Science in Mobile, AL
Escambia High School in Pensacola, FL
U.S.A. E-Week in Mobile, AL
Faith Academy in Mobile, AL
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VI.
Conclusion
It is apparent through this document that the USA Launch Society has made every effort possible to
prepare a detailed plan for the remainder of the competition. The payloads chosen for design include hazard
detection and wind turbine. The hazard detection payload will use a PixyCam to survey possible landing
hazards upon descent, and the wind turbine payload will determine the feasibility of generating a large enough
voltage to theoretically deploy the drogue parachute, as discussed in this document. The team has already
overcome numerous issues regarding the payload development including safety issues with alternative energy
supplies for deployments, a comparatively low budget and group size, timing issues for the wind turbine
power supply, etc. The team plans to continue the hard effort towards the payload development. Also, the
team is ecstatic about the exciting opportunities in the near future.
The proposed payloads offer increased safety and cleaner energy sources for the future of rocketry. In
rocky, rough climates such as Mars, hazard detection is a key to the future. With increased interest in future
missions to the Red Planet, innovations in hazard detection for descent vehicles is essential. This can save
NASA money and effectiveness. Due to questionable climate change conditions as well as financial concerns,
alternative energies are crucial. Alternative energies are the future of this world due to increasing restrictions
and fossil fuel depletion. Harnessing wind is an energy that will never cease to exist.
Acknowledgments
Before this report is complete, the USA Launch Society would like to thank the following people for their
support and expertise:
• Dr. Carlos Montalvo
• Dr. David Nelson
• Dr. John Steadman
• Dr. Dhanajay Tambe
• Dr. Richard Kramer
• Mr. John Hansel
• Mr. Kendall Brent
• Mr. Chris Short
All of these men have vastly helped to improve the efforts of the USA Launch Society. Their future
support is monumental to the longevity of this team and project.
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