Critical Design Review - University of South Alabama · PDF fileCritical Design Review University of South Alabama Launch Society Conner Denton, John Faulk, Nghia Huynh, ... electrical

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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American Institute of Aeronautics and Astronautics

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 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6

II Changes Since Preliminary Design Review 7

II.A PDR Feedback Questions and Answers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7

II.B Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7

IIIVehicle Criteria 8

III.AMission Statement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

III.B Success Criteria . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

III.C Major Milestone Schedule . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

III.DVehicle System Level Functional Requirements and Verification . . . . . . . . . . . . . . . . . 9

III.E Workmanship . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

III.F Additional Testing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

III.GRemaining Manufacturing and Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12

III.HDesign 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 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20

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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.KMission 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.NLaunch Concerns and Operations Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . 38

III.N.1 Final Assembly Checklist . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 40

III.OSafety 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 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 57

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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 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 63

V.C.2 Gantt Chart . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 66

V.D Educational Outreach . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 66

V.D.1 Purpose . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 66

V.D.2 Status of Outreach . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 67

V.E Schedule for Outreach Events . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 68

VI Conclusion 69

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I. Summary of Critical Design Report

I.A. Team Summary

I.A.1. 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 [email protected].

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. Launch Vehicle Summary

I.B.1. 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. Milestone Review Flysheet

I.C. 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. Changes Since Preliminary Design Review

II.A. 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

Critique Solution

1 Your apogee is listed as 6,000 ft. This is well aboveour FAA waver limit of 5,600 ft. Please prove to thereview panel that your apogee will not exceed 5,600ft. by CDR.

USA Launch Society has changed the vehicle’s motorbrand from Loki Research to Aerotech. With thatin mind, the team has changed the motor from aK960SF to a K550W motor, thus results in a lowersimulated apogee of approximately 5,160 feet.

2 Report mentions the use of cardboard as a construc-tion material. Can you confirm that this is a phenolicsubstance, and not card board?

The team has changed the vehicle’s body materialalong with the motor tube and centering rings to G-10 fiberglass. There will be brief stress and bucklinganalysis provided in the vehicle criteria ratio. III.I.5

3 The stability margin of 1.81 is low, and our chiefengineer would like to see a stability of 2.0 as therocket leaves the rail.

The team has an updated OpenRocket simulationthat results in a stability margin of 3.6 cal.

4 You have a high thrust to weight ratio. What arethe construction techniques for your fins?

Our thrust to weight ratio was initially miscalculatedin the PDR. The current thrust to weight ratio forthe new configuration for the vehicle is 6.15.

5 You have a slow descent velocity which can lead toexcessive drift. In general, descent velocities of 80-100 ft.

s and 17-20 ft.s for the drogue and main chutes

respectively will work for a rocket of your size andweight.

The team has taken into consideration the recom-mended descent velocities. Team members have per-formed analytical computations and simulations thatresults in a descent velocities of 131 ft.

s and 22.07 ft.s

for the drogue and main parachutes, respectively.

6 What are you generating electric current for? The team will be generating electric current to serveas a theoretical back-up voltage supply in order todeploy the drogue parachute. Data will be logged onan Arduino from the turbine to measure the voltageto ensure that the voltage is high enough to theo-retically ignite black powder to deploy the drogueparachute.

7 Where are you placing your turbines? The team has chosen for the turbine(s) to be placedin between the CP and CG to ensure a better stabil-ity of the rocket.

8 Does your turbine use an air scoop? If so, how doesit affect the stability of the rocket?

The team has designed air scoops to be placed overholes machined through the booster bay and turbinepayload. The scoops will be used to direct the airflow over the turbine. A scaled down version of therocket vehicle with scoops will be tested in the windtunnel at USA.

II.B. 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. Vehicle Criteria

III.A. 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.

Vehicle Requirement Requirement Features

1 The vehicle should deliver two

payloads to an altitude of 5,280

feet from the ground.

In order to achieve this altitude, the aerodynamics of the rocket will be maxi-

mized by selection of the ogive nose cone and the straight-tapered fins. Launch-

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-

mined using one, on-board baro-

metric altimeter. The rocket will

contain one altimeter for both

deployments and one predeter-

mined to be the scoring altime-

ter.

USA’s rocket will contain two barometric altimeters. One altimeter will control

both the main and drogue deployments while the other altimeter will serve as

the official scoring altimeter.

3 The vehicle must be reusable

and recoverable after each launch

with minimal or no damage.

The exterior of the launch vehicle will be coated in a selected material to

increase the strength of the body.

4 The launch vehicle shall have a

maximum of four independent

sections.

USA will have a two piece rocket. One section will carry the drogue deployment

and the turbine system while the second section will contain the main parachute

and the hazard detection camera.

5 The launch vehicle shall be lim-

ited to a single stage.

USA’s launch vehicle will perform its only motor ignition at takeoff.

6 The vehicle must be quickly pre-

pared for flight within two hours.

Prior to the launch date, the team will practice preparing the rocket to ensure

timely delivery.

7 The launch vehicle shall be capa-

ble of remaining in launch-ready

configuration at the pad for a

minimum of 1 hour without los-

ing the functionality of any crit-

ical on-board component.

When the team is practicing the timed vehicle assembly, they will also test to

ensure the rocket maintains functionality on the launch pad for a minimum of

one hour.

8 The launch vehicle shall be capa-

ble of being launched by a stan-

dard 12 volt direct current firing

system.

The launch vehicle will be capable of accommodating the 12 volt firing system

by utilizing e-match insertion into the motor.

9 The launch vehicle shall use a

commercially available solid mo-

tor propulsion system using am-

monium perchlorate composite

propellant.

All USA motors will be purchased from legitimate model rocketry vendors.

These motors will be NAR, TRA, and CAR certified.

10 The total impulse provided by a

launch vehicle shall not exceed

5,120 Newton-seconds (L-class).

Currently, USA plans to use a K-class motor which falls short of the 5,120

Newton-seconds total impulse.

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11 Pressure vessels on the vehicle

shall be approved by the RSO.

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.

Since there will be no pressure vessel on-board the launch vehicle, no extra

measures will be taken.

12 All teams shall successfully

launch and recover a subscale

model of their full-scale rocket

prior to the CDR.

The team plans to launch a subscale of the final rocket by mid-December.

13 All teams shall successfully

launch and recover their full-

scale rocket prior to the FRR in

its final flight configuration.

The full scale launch will be performed in early to mid-February to ensure

timely delivery of results by the FRR.

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 gener-

ation 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. Integrity of Design

III.I.1. 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

2ρV 2AsCD (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 swept-

tapered 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 =CTCR

(2)

III.I.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 = V ρ (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 Value

Fracture Strength 38000 psi

Flexural Modulus 2700000 psi

Poisson’s Ratio 0.12

Compressive Strength 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 deter-

mine 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

σcr =Eh

a√

3(1− ν2)(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 Value

Length (l) 94 inches

Radius (a) 2 inches

Thickness (h) 0.0625 inches

Critical stress (σcr) 49068.5 psi

Critical load (Fcr) 3904.75 lb.-f

III.J. 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 parachutedeployment 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 ft− 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 ft/s:

Ek =1

2mV 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 ft/s, the kinetic energies are as follows: Nosecone = 2.79 ft− lbf , P.A.P. =

36.26 ft− lbf , and the Booster = 52.71 ft− 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 ft/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 ft− lbf , P.A.P. = 173.62 ft− lbf , and the Booster = 252.37 ft− lbf .

The kinetic energy of each component at the deployment of the main parachute is determined using the

velocity at deployment of 131 ft/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 ft− lbf , P.A.P. = 1525.13 ft− lbf , and the Booster = 2216.34 ft− lbf .

III.J.4. Safety and Verification Plan of Recovery System

Recovery System Requirement Requirement Features

1 The launch vehicle shall stage

the deployment of its recov-

ery devices, where a drogue

parachute is deployed at apogee

and a main parachute is deployed

at a much lower altitude.

The launch vehicle will use an altimeter at apogee to deploy the drogue

parachute while a separate altimeter will be used to deploy the main parachute.

2 Teams must perform a success-

ful ground ejection test for both

the drogue and main parachutes

prior to the initial subscale and

full-scale launches.

These parachute testings will be done at the team’s regular launch field prior

to both of the aforementioned flights in order to prove recovery system effec-

tiveness.

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3 At landing, each independent

section of the launch vehicle shall

have a maximum kinetic energy

of 75 ft-lbf.

In order to ensure that the independent sections do not exceed the maximum

kinetic energy, the team will model the descent of the rocket in OpenRocket

where the velocity of the rocket at impact with the ground will be calculated.

Then, the kinetic energy of each section will be calculated using the equation

E = 12mV

2.

4 The recovery system electrical

circuits shall be completely inde-

pendent of any payload electrical

circuits.

Each parachute deployment will be controlled by its own altimeter. Each pay-

load will be controlled by electrical systems that are unrelated to the recovery

system altimeters.

5 The recovery system shall con-

tain redundant, commercially

available altimeters.

The launch vehicle will contain two PerfectFlite StratoLoggers and one Tele-

Mega, all of which are commercially available.

6 Motor ejection is not a permissi-

ble form of primary or secondary

deployment. An electronic form

of ejection must be used for de-

ployment purposes.

Each StratoLogger will initiate a black powder ejection charge for their respec-

tive parachute deployments at apogee and low-altitude main deployment.

7 A dedicated arming switch shall

arm each altimeter, which is ac-

cessible from the exterior of the

rocket airframe when the rocket

is in the launch configuration on

the launch pad.

USA will use either simple wire twisting switches or small flip switches to

initiate the altimeters from the exterior of the rocket.

8 Each altimeter shall have a ded-

icated power supply.

Each altimeter or GPS will have its own 9 volt battery.

9 Each arming switch shall be ca-

pable of being locked in the

”ON” position for launch.

If twisting wires is chosen, they will be tightly secured in a specific hooked

formation. If the flip switches are used, they will automatically be locked into

position once they are activated.

10 Removable shear pins shall be

used for both the main parachute

compartment and the drogue

parachute compartment.

Nylon shear pins will be used to attach the components. The exact properties

of these shear pins will be determined by the CDR to ensure proper shear

calculations.

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11 An electronic tracking device

shall be installed in the launch

vehicle and shall transmit the

position of the tethered vehicle

or any independent section to a

ground receiver.

The exact tracking device is yet to be determined at this time. USA’s launch

vehicle will remain entirely tethered together, therefore, there is no need for

multiple transmission devices.

12 The recovery system electronics

shall not be adversely affected by

any other on-board electronic de-

vices during flight.

The two recovery altimeters will remain in an independent bay from the elec-

tronic tracking device in order to avoid inadvertent excitation of the recovery

system.

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 func-

tioning 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 fts2 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. Drag

III.L.1. 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

2CDρV

2A (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 +

1.5

(L/d)1.5

]SWSBT

(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

Cf =1.33√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

[ln(0.06Re)]2− 1520

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 =0.029√

CDN + CDBT(13)

III.L.3. 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 = CDOFCRSBT

d

2xnumberoffins (14)

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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 = 1.2SLLSBT

(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 = πdoutl + πdinl (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

~Mθ = CmqSdj1 (18)

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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 =2

Sd[lAb − V olume]

sinα

α(19)

III.L.6. 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

V =

√8mg

πρCDD2(20)

In the above equation, m is 0.45067 slugs, g is gravity, ρ is 0.0023769 slugs/ft3, CD is 0.8, and D is the

diameter of the parachute which is 6.31758 ft. Inserting these values into the equation, the speed of descent

is determined to be 22.07 ft/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 ft/s (10 mph = 14.67 ft/s). This

provides a result of 265.82 ft at 5 mph, 531.64 ft at 10 mph, 797.46 ft at 15 mph, and 1063.28 ft 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 successfullaunch.

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. Safety and Environment (Vehicle and Payload)

III.O.1. 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 con-

sequently 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 ex-

pected 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 dam-

age 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 Payload Requirement Subsystems Analysis

Pixy Camera (CMUcam5) X, Y, and Z coordinatesfor hazard detection

Arduino Uno, brightly col-ored tarps, 9V battery,Stratologger altimeter

Identify the exact locationby association with theheight given by the altime-ter

DC Motor Turbine Create at least 2V duringhalf the time to apogee

Arduino Uno and 9V Bat-tery

Provide enough chargefrom the turbine to theo-retically power an e-matchat apogee

IV..1. PixyCam

Table 24: Camera selection criteria.

Camera Module Weight Dimensions Cost Application

Pixy Camera (CMUcam5) 0.1 lb 2.2 x 1.5 x 2.2 in $69.00 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

(xγsin(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

(yγsin(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. Payload Testing

IV.A.1. 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.

Test 1 Test 2

xactual (in) 12 1

yactual (in) 29 16

zactual (in) 228 34

xγ 14 6

yβ 20 80

xα 150 150

yβ 100 100

α 37.5 37.5

β 37.5 37.5

xcalculated (in) 12.975 0.8217

ycalculated (in) 18 11.671

IV.A.2. DC Turbine

Testing will be performed prior to launch to provide feasibility for the payload. First we will test the turbine

voltage output in a wind tunnel set to the highest speed of launch ascent (678 ftsec ) 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. Uniqueness and Difficulty of Payloads

IV.B.1. 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).

Vc = Vs

((1− e

−tRC )

)(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.

Vc = Vse−tRC (26)

The e-match consists of a black powder charge. Black powder charge has a flash of approximately 801-

867 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(27)

The resistance found is 1.1236 Ω. Resistance of 30 gage copper wire was found to be 0.339 Ωm ; thus with

4 inches of copper wire used the resistance is 0.034442 Ω. Resistance of nichrome wire was found to be 6.75

Ωft . 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

Rtotln(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. Payload Concept Features and Definition

IV.D.1. 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. Project Plan

V.A. Budget Plan

V.A.1. 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 Price Reasoning

Airframe $120 The team is budgeting $120 for the airframe including

the nose cone.

Fins $28 $28 will be budgeted for the construction of the fins.

Motor Set $113 Complete motor - includes one casing, one nozzle, one

forward bulkhead, two snap rings, and one nozzle washer.

Center Rings $20 $20 will be used for the center rings.

Miscellaneous $100 The team is adding $100 for miscellaneous items possibly

needed for the airframe.

Subtotal $381 $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 Price Reasoning

E-matches $10 There will be a set $10 budget for e-matches.

Cable Cutters $55 Used for dual deployment purposes from the same bay.

Main Parachute $225 The team will manufacture the main parachute with

the given budget.

Drogue Parachute $6 The team will manufacture the drogue parachute with

the given budget.

Black Powder $20 There will be a budget of $20 for black powder.

Shock Cord $10 Minimal cost due to the team having an ample amount of

shock cords from previous launches.

Miscellaneous $100 Flexible amount in the event of emergency.

Sewing Machine $0 Members of the team possess a sewing machine.

Subtotal $426 The projected budget strictly for recovery components.

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Table 28: Altimeter bay budgeting.

Altimeter Bay Price Reasoning

Altimeter(x2) $140 The team will be using two Perfectflite Stratologgers altimeter for the vehicle.

Shear Pins(x2) $8 Both 2-56 and 4-40 shear pins are available in the Aerospace

Propulsion Lab.

Misc. Components $100 Extra parts and supplies such as switches, screws, bolts,

batteries, glue, etc.

E-match Material $100 Includes prefabricated wires or wires and ignition.

Bay Casing $70 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 Charge $130 Ejection charging material needed to separate the sections

of the rocket includes black powder, pyrodex, and CO2 kits.

Subtotal $548 The subtotal describes the range of price for the components

and parts for the altimeter bay to be between $548.

Table 29: Certification budgeting.

Certifications Price Reasoning

Certification Kits (x2) $200 The team will budget $200 for certification kits

to be flown September.

Certification Motors (x2) $160 The team is budgeting $160 for certification motors.

Subtotal $360 The total cost of the certification is $360.

V.A.2. 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.

Outreach Price Reasoning

Model Parachutes (x10) $75 The team is budgeting $75 that is divided into 4 yards

of nylon and lines for the parachutes.

Straws $10 $10 is budgeted to buy regular straws for a

pressure rocket system.

Gas per Visit (x7) $100 $100 is designated for gas to attend the outreach

activities.

Rocket Motors (x10) $200 $200 is budgeted for the mini rockets that will

be launched at schools.

Subtotal $385 The subtotal cost for the the outreach activities is $385.

V.A.3. 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.

Travel to Samson, AL Price Reasoning

Samson, AL (Driving x10) $900 The team is projected to attend a potential of 10

available launches. One will be held once every 3

weeks starting September 12, 2015.

Subtotal $900 Total for transportation to launches in Samson, AL.

V.A.4. 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 Price Reasoning

Driving (3 cars) $405 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.

Food (x5) $300 $300 is budgeted for the teams food expenses.

Hotel (3 nights) $700 The team is budgeting $700 for hotel expenses for

approximately 8 people.

Subtotal $1405 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 Price

Rocket Structure $381

Payload $400

Recovery Components $426

Altimeter Bay $548

Certifications $360

Outreach $385

Travel to Samson, AL $900

Travel to Huntsville, AL $1405

Total Project Cost $4806

V.B. 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.


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.


9-20 PDR video teleconferences.

December 2015

4 Critical Design Review (CDR) Q&A.


January 2016

15 CDR reports, presentation slides, and flysheet posted on the

team website by 8:00 A.M. Central Time.


19-29 CDR video teleconferences.

February 2016

3 Flight Readiness Review (FRR) Q&A.



March 2016

14 FRR reports, presentation slides, and flysheet posted to team

website by 8:00 A.M. Central Time.


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. Educational Outreach

V.D.1. 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 studentsat 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 thelaunch rod for the kit rocket.

V.E. Schedule for Outreach Events

Table 35: Outreach events schedule.

Date Location of Outreach

November 13, 2015 T.R. Miller High School in Brewton, AL

December 14, 2015 W.S. Neal Elementary School in Brewton, AL

January 19, 2016 Alabama School of Math and Science in Mobile, AL

February 11, 2016 Escambia High School in Pensacola, FL

February 15-19, 2016 U.S.A. E-Week in Mobile, AL

February 26, 2015 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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References

1National Aeronautics and Space Administration. Nasa student launch non-academic handbook. pages 58–59, 2015.

2Tripoli Rocketry Association. Tripoli level 2 certification. http://www.tripoli.org/Level2. Accessed: 10/24/2015.

3Tripoli Rocketry Association. Code for high power rocketry. page 1, 2012.

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