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FLIGHT READINESSREVIEW (FRR)Charger Rocket WorksUniversity of Alabama in HuntsvilleNASA Student Launch 2013-14Kenneth LeBlanc (Project Lead)Brian Roy (Safety Officer)Chris Spalding (Design Lead)Chad O’Brien (Analysis Lead)Wesley Cobb (Payload Lead)

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Prometheus Flight Overview

Payloads Here

Payload DescriptionNanolaunch 1200 Record flight data for aerodynamic coefficients

Dielectrophoresis Use high voltage to move fluid away from container walls

LHDS Detect and transmit live data regarding landing hazards

Supersonic Coatings Test paint and temperature tape at supersonic speeds

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Technology Readiness Level

http://web.archive.org/web/20051206035043/http://as.nasa.gov/aboutus/trl-introduction.html

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Outreach• Adaptable for different ages and

lengths• Supporting activity

• Water Rockets• Completed

• Science Olympiad• Challenger Elementary• Discovery Middle• Horizon Elementary

• Numbers• Education Direct: 466• Outreach Direct and Indirect: 723

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On Pad Cost

Payload

Hardware

Recovery

Propulsion

$- $200.00 $400.00 $600.00 $800.00 $1,000.00 $1,200.00 $1,400.00

$976.98

$1,212.19

$204.53

$823.91

Total Rocket Cost: $3,217

Cost

Syst

em

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DESIGNTeam Members:Chris Spalding - Team Lead

Andrew Mills - Prototype AssemblerJordan Lee - Designer

Josh Thorne - Coatings David Zaborski - Recovery Designer

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Hardware Changes:

1. Printed Nylon2. Transition coupler to

accommodate nose cone mold error

3. Flat bulk head and additional coupler joint

4. Flat bulk head 5. ABS plastic Brackets secured

with Chicago screws

Design Details:

• 34lbs

• 40Gs acceleration

• Geometric similarity to NASA

Nanolaunch prototype

• Nanolaunch team requested

maximum use of SLS printed

aluminum

1 2 3 45

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Interfaces (1)

# Component Interface Load Locations

1 Pitot Probe Threaded insert epoxied in and secured to Nose cone shaft

Tension from pitot shaft, compression from nose cone, aerodynamic forces

2 Nosecone Slip fit with shear pins Compression from pitot probe and slip

3 Nosecone Payload Threaded to nosecone shaft Acceleration forces, passed through nose cone shaft

4 Nosecone shaft Threaded to pitot probe insert Tension loads between the nose cone bulkhead and pitot probe, compression/tensions from payload acceleration forces

5 Nosecone bulk head slipped over payload shaft Tension from payload shaft ring nut.

6 Nosecone shaft nut Threaded to nosecone shaft Tension from payload shaft

7 Recovery Package Shock cord, quick links, ring nut

Tension from ring nuts, aerodynamic forces

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Interfaces (2)

# Component Interface Load Locations

8 Payload Slipped onto payload shaft, constrained between nuts and bulkhead

Acceleration forces, passed through payload shaft

9 Lower Coupler Epoxied to lower body tube, in compression between body tube sections with payload shaft

Compression from middle and lower body tubes, aerodynamic forces

10 Payload Shaft Threaded to motor case, lower bulkhead, and ring nut

Tension between bulkheads and ring nut, compressive and tensile forces from payloads under acceleration

DELETE ROW

12 Motor case Threaded to payload shaft Outside manufacture; loaded in designed manner

13 Fins /Fin Brackets Bolted to lower body tube, T-nuts inside body tube.

Aerodynamic acceleration forces, resulting tension from body tube.

14 Thrust Ring Held in compression between motor case and body tube.

Compression from motor case

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Thrust Ring

• Machined 5086 Aluminum

• Will be Analyzed with FEA

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Fin Assemblies

• ABS has been proof loaded to 75 lbs

• 3D printed Laser sintered nylon brackets have been

ordered

• Bolted to body

• Binding post fin attachment

Currently have sets of fin brackets in abs plastic and fiber reinforced nylon.

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Body Tube

• Three body tube pieces joined

with nylon printed couplers

• Carbon composite

• FEA, destructive testing and

hand calculations done to assess

strength

• Large margin of safety and low

weight

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Payload Shaft

• 7075-T6 Aluminum threaded shaft 3/8-16

• Preloaded in tension

• FEA and hand calculations show significantly over strength requirements

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Payload Shaft Load Paths

• Carries thrust loads into payloads and recovery forces into lower rocket, as well

as providing assembly method for payloads, body tubes and recovery harness

• Red Arrow indicates motor loads from thrust ring through body tube

• Green arrow indicates motor loads passed through payloads

• Blue arrow indicates recovery forces passed through payload shaft

• Orange arrow indicates motor case retention force

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Coupler Rings

• Sintered nylon (potentional to be

reinforced with aluminum or

carbon fiber)

• Aft coupler retained by payload

shaft preload. Also, one side will

be epoxied to the body tube.

• Fore coupler retained by nose

cone shaft and shear pins

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Nose Cone Assembly

• All components retained by shaft similar to payload shaft

• Carbon fiber nose cone shroud and bulkhead

• Bulkhead is secured with tension from the nose cone payload shaft (seen

on next slide)

• Contains pitot pressure and accelerometer/ gyro data package

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Nose Cone Assembly

• Coupler is designed for slip fit and

secured with shear pins. Secured with

tension in the payload rod.

• The new design allotted more space

for the recovery system

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Pitot Probe

• Allows measurement of static pressure

along with supersonic AND subsonic total

pressure

• Unique and original design which could

only be made with 3D printing techniques

• Helps fulfill our Nanolaunch request to

explore selective laser sintering in

original ways.

Old Design

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Pitot Probe

• Secured with threaded insert

epoxied into center (blue

part)

• Connection ports are now

open to attachment by

epoxying tube directly

• The change allowed for

simplified 3D printing

Manufactured out of glass reinforced nylon.

New Design

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Vehicle Success CriteriaRequirement Criteria Verification

Safe launch No harm to anyone or the rocket

Safety analysis before launch. No harm to anyone or rocket

Recoverable and Reusable No Damage to the rocket or payloads

Visual inspection of structures for verification post launch

Geometric similarity to the Nanolaunch 1200 prototype

Design Vehicle with High fidelity to Nanolaunch Project Geometry

Rocket matches scaled design of Nanolaunch during fabrication

Supersonic Flight Reach Mach 1.4 Review data from accelerometers and pitot pressure sensors post launch back at the lab

Vehicle must be assembled and ready to fly in reasonable time

Vehicle must be assembled in less than 3 hours from arrival at launch field

Practice procedures to get team fluent in the assembly

Payloads must be integrated into vehicle design.

Payloads must be receive and send data from ground stations

Design accommodates for necessary communications and payload operations pre launch

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ANALYSISTeam Members:

Chad O’Brien - Team Lead

Sarah Sheldon - Design Analyst Garrett Holmes - LHDS AnalystTryston Gilbert - Trajectory Analyst

Fernando Duarte - Prototype Design Analyst

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GENERAL ROCKET MISSION PERFORMANCE CRITERIA

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General Rocket Flight Performance CriteriaRequirement Success Critera Verification

Safe Launch Operations Vehicle Launch does not cause injury to launch crew or bystanders.

Vehicle maintains safe heading and travel.

Aerodynamic Stability for launch Launch velocity must exceed minimum speed for stable flight

Visual observation of vehicle behavior off the rail coupled with empirical data from sensors.

Transonic flight data Stay within Mach 0.7-1.4 and collect usable data

Receive readable data from accelerometers and pitot pressure sensors

Achieve Sustained Transonic Flight Stay within 750 ft/s – 1400 ft/s

Data from accelerometers or pitot pressure sensors

Supersonic flight reached Reach above Mach 1 Data from accelerometers or pitot pressure sensors

Fin Design Supports Supersonic Flight Fins should be ready to fly after short post flight inspection and new flight preparations without modification.

Visual inspection and simple sturdiness test to ensure fixtures and material are ok to fly.

Meets Drift Requirement Lands within 5000 feet of the launch tower Tracker and GPS will be used to verify position on landing and a distance from launch sight will be calculated.

Safe Ground Energy Impact Levels Components must impact with less than 75 feet pound force

Review of flight data to see impact speed.

Recoverable and Reusable Rocket can be launched again without significant alteration

Rocket can be launched again in same day

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FLIGHT SIMULATIONS

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Mass StatementSubsystem Mass (lbs)

Payload 2.93

Recovery 4.7

Airframe 12

Motor Case 7.0

Propellant 7.4

Total Dry System 27

Total Wet System 34

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Prometheus Simulation• RockSim Software Package• Motors

• Primary: CTI4770 – 98mm• Secondary: AeroTech K1499 – 75mm

• Estimated Dry Mass at 27 lbs• Launch Conditions for Salt Lake City

• ASL – 4210 feet• Temperature – 72 ˚F

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Final Motor Selection - CTI M4770-P

• ISP – 208.3s• Loaded Weight: 14.4lb• Propellant Weight: 7.3 lb• Max Thrust: 1362 lbf

Prometheus’ Static Margin

• Launch Static Margin – 1.7• Burnout Margin – 4.5

CG at 85.8”

CP at 93.6”

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Prometheus Simulation

• Max Altitude – 15,700 feet• Max Velocity – 1600 feet per second• Max Acceleration – 40 Gs

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Prometheus’ Static Margin• Pre-Launch Static Margin: 1.7• Burnout Static Margin: 4.5

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Monte Carlo Analysis

Altitude:

Mach:

Accleration:

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Drift Analysis

• 500 Cases for each cross wind. • High probability of landing within the 5000 foot requirement

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Variation in Flight Time

• Time variance directlyaffects the radial landingdistance.

• Dependent on high speedcoefficient of drag for drogue

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Plan B Motor: Aerotech K1499

• Altitude – 2100 feet• Velocity – Mach 0.25• Acceleration – 16 G’s

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FIN FLUTTER ANALYSIS

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The Equations for Fin velocity• t = thickness of fin• AR = aspect ratio • l = taper ratio• G = shear modulus.• C = root chord• P = air pressure• a = speed of sound

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The Equations for Fin Velocity • S - Wing Area• b - Semi-span• Cr - root chord • Ct - tip chord• T - Temperature of air

• Area = 0.5(Ct + Cr)b• AR = b2/S• l = Ct/Cr

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Prometheus FinGiven: Variable Value Units

Cr 8.31 in.

Ct 4.75 in.

t 0.17 in.

b 4 in.

G* 5.00E5 psi

S 26.12 sq. in.

AR 0.61 dimensionless

l 0.57 dimensionless

h 3000 ft

T 48.32 F

P 13.19 psi

V 2071 ft/s

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Assumptions• Shear Modulus: 5E5 psi• Isotropic Layup• Applied Max Velocity of 2000 ft/s• Solved for Material Thickness

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Conclusion• At exactly t = 0.17 inches, max V = 2071 ft/s • Designed Max V = 2000 ft/s• Projected Max V = 1600 ft/s• The safety range is accounted for with current design

and material of Prometheus

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Buckling Analysis

• Used Euler’s Buckling Equations

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Recovery System

• Single Separation Point• Main Parachute

• Hemispherical• 12 ft • Cd 1.3 ( flight test)• Nylon

• Drogue Parachute• Conic• 2.5 ft• Cd 1.6 (flight test)• Nylon

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Deployment Bag• Nomex Fabric• Kevlar Thread• Fiberglass Rod Inserts for Rigidity• Shroud line “daisy chained” and coiled in bag section.

Bag Section

Fiberglass Rod inserts

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Main Parachute

• 12 Feet Semi-Hemispherical• Ripstop Nylon• Custom Seam • 14 Gores• Shroud Lines: 0.125in x 550lb Paracord

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Sewing Technique• Multi Method Gore Stitch• Straight stitch• Zigzag stitch• Biased Tape Reinforced Joints• Edges hemmed using serge roll.• Joints Reinforced with Nylon Straps.

Seam Cross Section

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Construction MaterialsPart Material

Main Canopy Ripstop Nylon

Thread Polyester/Kevlar

Line Anchor Points 0.019” thick Nylon

Swivels 316 SS

Eyenut Steel

¼” and 3/8” nuts Steel

¼” and 3/8” washers Steel

Quick links 316 SS

Shroud Lines 550 Parachord

Main Shock Chord ¼” diameter Kevlar

Deployment bag Nomex

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Recovery System Deployment Process

• Stage 1• 2 seconds after apogee • nose cone separates• release the drogue

• Stage 2• The Tender Desenders

release• Stage 3

• Main parachute falls out deployment bag/burrito

Eye bolt

L.H.D.S

Tethers

Black Powder Charge

Drogue

Main Parachute InDeployment bag/Burrito

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Deployment Process

Stage 1: Drogue Deployment Stage 2: Tether Separation Stage 3: Final Decent

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GPS Tracking• GPS Module: Antenova M10382-Al

• GPS lock from satellites• Transmits data through XBee RF module• 8 ft accuracy with 50% CEP (Circular Error Probable)• 3.3 VDC supply voltage• 22 to 52 mA current draw

• Since CDR, redundant GPS Unit: “Tagg Pet Tracker” no longer included

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Radio Transmission• RF Module: XBee-PRO XSC S3B

• 900 MHz transmit frequency• 20 Kbps data rate• 9 mile LoS range• 250 mW transmit power• 3.3 VDC supply voltage• 215 mA current draw• 1.5+ hr battery life at max sensor sample rate• Laptop ground station

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GPS/RF Module Ground Testing• Stationary ground station• Transmitter driven away from receiver, increasing LoS obstructions

• Obstructions were increased until data dropout

• Test was a success in worst case scenario terrain conditions

Data Dropout

LoS

Ground Station

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GPS/RF Module Flight Testing• Full-scale test on April 12, 2014

• Successful deployment of module after apogee• Failure to transmit/receive live GPS data

• Suspected causes include• Pre-flight damage to Antenova M10382-Al GPS chip• Failure to preform pre-flight testing• Sustained damage from crosswinds on landing

• Mitigation of future failure• Inactive device management• Testing added to pre-flight SOP• Comprehensive shielding on payload sled

• Further flight testing scheduled before competition

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Energy and Velocity at Key Points

Wind Speed Range (MPH) Average Drift (feet)3-4 1388

8-14 3358

15-25 5962

Stage of Recovery

Altitude (ft) Velocity (ft/s)

Energy (ft*lb)

1 15190.2 50.175 878.5

2 1000 98.58 3391.23

3 0 9.702 32.87

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TESTING ANDVERIFICATIONBrian Roy – Safety Officer

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Testing ProceduresTest

Requirement Identified

Develop Operating

Procedures

Review of Procedures

by PRC Staff

Procedure Approval by

PRC Director

Identify Red Team

Members for Test

Review of Operating

Procedure with Red Team

Approval of Red Team Members

Testing

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Subscale Testing and ResultsSub-Scale Flight Test Matrix

Type of Test Test Goals Results

Sub-Scale Flights Verify the vehicle stability margin and flight characteristics. Successful (2/8/14)

Flight ElectronicsEnsure that payload records proper data and that launch detect functions properly.

Successful (3/8/14)

Recovery System Hardware

Test hardware that will allow for a single separation dual deploy setup in full-scale vehicle.

Successful (4/12/14)

Parachute DesignVerify construction techniques are adequate and determine effective drag coefficient.

Successful (2/22/14)

High Acceleration Flight (40+ G’s)

Ensure that avionics will survive launch forces of full-scale. Successful (3/8/14)

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Sub-Scale Flight #1• Goals: Verify stability of Prometheus’ outer profile.• Test Date: February 8, 2014. Childersburg, AL.• Vehicle Configuration: Arcas HV kit with additional body

tube sections to obtain proper outer profile and Nanolaunch payload to collect data flown on I-205 motor.

• Flight Results: Successful flight and recovery. Payload failed to activate, no data collected.

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Subscale Flight #1 Data

• Apogee: 1,573 feet AGL.• Max Velocity: 279 ft/s.• Time of Flight: 63.9 seconds.• Recorded Using a PerfectFlite SL100

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Subscale Flight #2• Goals: Verify proposed recovery system design and

retest Nanolaunch payload.• Test Date: February 22, 2014. Manchester, TN.• Vehicle Configuration: First subscale vehicle with

revised fin design and in-house manufactured drogue parachute. Flown on an Aerotech I-600R.

• Flight Results: Main parachute did not deploy. Nanolaunch payload prematurely triggered, data not usable.

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Subscale Flight #2 Data

• Apogee: 4,156 feet AGL.• Max Velocity: 597 ft/s.• Time of Flight: 128.6 seconds.• Recorded Using a PerfectFlite SL100

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Subscale Flight #3• Goals: Verify proposed recovery system design, subject

electronics to high G-loads.• Test Date: March 8, 2014. Childersburg, AL.• Vehicle Configuration: Second subscale vehicle with

CTI J-1520 V-max.• Flight Results: Main parachute became tangled in shock

chord, failed to deploy. Data successfully collected by Nanolaunch payload. No adverse effects due to G-loading (~25 G’s).

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Subscale Flight #3 Data

• Apogee: 7,758 feet AGL.• Max Velocity: 1,208 ft/s.• Time of Flight: 210 seconds.• Recorded using a PerfectFlite SL100

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Prototype Flight #1• Goals: Test full-scale recovery system,

LHDS, in-house manufactured parachutes, and 3-D printed parts.

• Test Date: April 12, 2014. Manchester, TN.• Vehicle Configuration: 5.5” diameter

fiberglass rocket with simulated 4.5” recovery bay and full-scale payload retention system. Flown on an Aerotech K-1499.

• Flight Results: Successful flight and recovery of all components. GPS lock not obtained due to short flight time.

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Prototype Flight #1 Data

• Apogee: 1,259 feet AGL.• Max Velocity: 279 ft/s.• Time of Flight: 65.5 seconds.• Recorded using a PerfectFlite SL100

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PAYLOADSTeam Members:

Wesley Cobb - Team Lead

Bronsen Edmonds - Sensor DeveloperTyler Cunningham - Dielectrophoresis

Shawn Betts - LHDS Samuel Winchester - Tracking

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Payload Integration

LHDS

Dielectrophoresis and Aerodynamic

Coefficient Payload

Aerodynamic Coefficient Payload

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Nanolaunch Experiment Overview• Calculating Aerodynamic Coefficients

• Pitching moment Coefficient• Drag Coefficient• Measure base pressure

• Two separate sensor packages• Accelerometers• Gyroscopes• Pressure sensors

• Similar not identical• Nosecone

• Pitot probe• 60 PSI• 100 PSI

• Near CG• Base pressure sensors

• 30 PSI• Designed for future use

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Nanolaunch Rocket Rotation Results• Ground Test Rocket Spin Curve Fit

• Uses Rot-y axis from Gyro• Fit Minimizes R^2 Value for Exponential Decay Sine Wave• Ground Test Indicates 1.1709 Hz

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Nanolaunch Rocket Rotation Verification• Verified using FFT• First peak -> Due to offset• Low frequency = 1.2 Hz• Fast frequency = 11Hz (Low amplitude noise)

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High G Accelerometer Data(Ground Test)

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CG Transducer Pressure Data(Used for Calibration) Setra 830E

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CG Pressure Sensors Calibration Curves

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Transducer Uncertainty Results

• Ways to decrease uncertainty:• Optimize gain resistor value• Measure more data points in the Vacuum region to improve calibration curve• Purchase Higher Precision Transducer

Pressure (psig)

Regression Uncertainty

(psi)

Altitude Uncertainty

(ft)

Confidence Interval X ±

U (ft)

0 0.2080 386.3 X ± 386.3

0.5 0.2050 380.5 X ± 380.5

5.5 0.2070 384.3 X ± 384.3

6.5 0.2100 390 X ± 390

10.5 0.2350 438 X ± 438

5.5 0.2070 384.3 X ± 384.3

0.5 0.2050 380.5 X ± 380.5

0 0.2080 386.3 X ± 386.3

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Nose Cone Shock Interactions for Pressure Sensor Consideration

• Bow Shock Due to Blunt Tip• Measure Stagnation Pressure and Static Pressure After the Shock

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Converting Pitot Probe Data to Velocity• Using Normal Shock and Isentropic Relations• Po2/P1 ratio can be directly looked up to find M1 and M2

Atmospheric Pressure Measured Data Calculated Ratios Before Shock After Shock

P1 [Pa] Po2 [Pa] P2 [Pa] Po2/P1 P2/P1 M1 M2

101325 101353.4 101325 1.000 1.000 0.020 0.020

101325 102036 101325 1.007 1.000 0.100 0.100

101325 104190.6 101325 1.028 1.000 0.200 0.200

101325 107853.4 101325 1.064 1.000 0.300 0.300

101325 113134.6 101325 1.117 1.000 0.400 0.400

101325 120193 101325 1.186 1.000 0.500 0.500

101325 129240.4 101325 1.276 1.000 0.600 0.600

101325 140548 101325 1.387 1.000 0.700 0.700

101325 154453.8 101325 1.524 1.000 0.800 0.800

101325 171371.3 101325 1.691 1.000 0.900 0.900

101325 191801 101325 1.893 1.000 1.000 1.000

101325 216110.2 126150 2.133 1.245 1.100 0.912

101325 243939 153305 2.407 1.513 1.200 0.842

101325 274952.5 182892 2.714 1.805 1.300 0.786

101325 308964.5 214809 3.049 2.120 1.400 0.740

101325 345849 249057 3.413 2.458 1.500 0.701

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Velocity Verification• Ready to calculate when full scale Po2 and P2 are

measured by the Pitot probe. • The Mach vs the ratio between the two measurements will

look like the following using normal shock relations:

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Nanolaunch Success Criteria• Objectives: Meet Team/NASA SLI Requirements and

Verify Those Were MetRequirement Success Criteria Verification

Velocity Verification Measure Pitot static pressure at the nose to

calculate Mach

Recover pressure data from the Pitot static

probes

Determine Axial Force Measure axial acceleration

Recover acceleration data in the axial

direction

Determine Angle of Attack

Measure gyroscope data at CG and the

nose to get Yaw, Pitch, and Roll

Recover gyroscope data from both

Beaglebone modules

Recoverable and reusable

Recover the payload and reuse it

Recover the payloads and be able to

relaunch again in the same day

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Outcomes and Nanolaunch Path Forward

• Outcomes:• Successfully Extracting Data• Calculated Rocket Spin• Pressure Sensors Calibrated• Payloads Fabricated and PCBs mounted

• Path Forward• Record More Launch Data for Data Comparison• Calibration of Gyro, High G Accelerometer• Find Higher Precision Transducers with Accuracy of +- 0.1% FFS

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Dielectrophoresis (DEP)• Fluid manipulation

• Electric field• Peanut oil

• Voltage• Voltage squared drives strength of electric field

• Fluid• Dielectric constant determines fluid interaction

• Electrode geometry• Gradient of electric field depends on geometry

Uniform Electric Field

Positive Region Negative Region

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Experimental Changes

• Electrode configuration: from parallel electrodes, to annular electrodes

• Voltage increase from 7kV to ~12kV

2012-2013 Configuration

2013-2014 Configuration

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DEP Testing• EMI Testing

• Test next to flight ready recovery system• Minus gunpowder

• Test next to Nanolaunch• Test and Prove design

• Test revised circuit• Structure tests

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DEP HV Output TestThis is the voltage probe used to test the Dielectrophoresis HV supply.

The readout from the probe showed that the HV supply was putting out 60kV

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DEP Success Criteria Requirement Success Criteria Verification

Microgravity environment

Reach apogee of flight to experience

microgravity environment

Retrieve accelerometer data to determine duration of

microgravity environment

Manipulate fluid with electric field

Noticeable collection of fluid around central

electrode

Retrieve camera and accelerometer data

Perform experiment without interfering

with other payloads

Reliable data collection from all

payloads adjacent to DEP

Rigorous preflight testing. Post flight analysis of data.

Recoverable and reusable

Fluid containers intact. No electrical shorts. Functional

electronics

Recover the payload. Return to flight ready state with no repairs

needed.

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Supersonic Paints and CoatingsUrethane Epoxy

Epoxy

•Urethane • Excellent retention• Abrasion resistant • Smooth Coating

•Epoxy Primer• Low film build• Excellent adhesion• Rough Coating

•Thermal tape• 3-5 second reaction time• Changes color at specific

temperatures• Excellent Adhesion

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SPC Testing• Oven Testing for Temperature tape

• Calibration of tape• Temperature sensitivity • Reaction time

• Flight Test• Subscale Test Flight• Full scale test launch

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Success Criteria of Paints and Coatings

Requirement Success Criteria VerificationEven film thickness Coverage of the

coatings is even and adheres correctly

Check for any defects post flight

Low coating weight Adds minimal weight to the rocket

Weighing the rocket before and after

applicationHigh heat resistant Coating unscathed

from thermal loadsNo discoloring of the coatings post flight

Recoverable and Reusable

Recover the payload and reuse it

Recover the payloads and be able to

relaunch again in the same day

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Landing Hazard Detection• Beaglebone

• NX-3000 USB Camera• Python Libraries• Established knowledge base

• 3 Methods of Analysis• Color detection• Edge detection• Shadow analysis

• Orientation• Use accelerometer to filter images of the ground

Full Scale LHDSRadio Shown

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LHDS Testing• Test Flights

• Full scale only• Alter method for different launch field

• Bench Test• White wall simulates salt flats• Colored paper as “hazards”• Google Map images

Hara Launch FieldManchester, TN

After Edge Detection

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Hazard Detection method• Original image

taken from Google Maps of Hara Launch site

• This image is analyzed by the Beaglebone searching for a range of green pixels.

• Green pixels are turned white

• All other pixels are black.

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Hazard Detection method• The original image

overlaid to test inspect if green images were flagged

• Then, a Canny Edge detection algorithm is run on the image to search for edges in picture.

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LHDS Success criteriaRequirement Success Critera Verification

Recoverable and Reusable If the Payload can be removed and replaced between subsequent flights

There are no missing screws or bolts or tools necessary to fix the LHDS.

Sustainable The mass simulators or payload does not hinder the Rocket’s takeoff or landing

If the payload remains inside its design, and the Rocket launch and landing are successful.

Non-damaging If no damage to the rocket is done by the mass simulators

If any scratches or dents are visible inside or outside the rocket near the payload

Communicable If the RF antenna can communicate successfully without impedance from the design

If a signal is reached from the ground base.

Camera Visibility If the camera can take pictures of the ground without any obstruction from the Rocket body after deployment

If the pictures have desired amount of ground-landing in them to be able to verify landing hazards

Communicable GPS RF module communicates with GPS module and ground

If GPS location is communicated to ground

Functioning Electronics If electronics are functional and record data properly

If data is recorded and readable for analysis

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Beaglebone Cape• Components

• ADLX 377 (Analog Accelerometer)

• L3GD20 (Gyroscope)• ADLX 345 (Digital

Accelerometer)• IC^2 connection to other

boards• Purpose

• Detect launch and initiate Data Acquisition

• Take measurements

Final Product printed by OHS Park

CAD Drawing

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GPS Antenna PCB• Components

• Xbee Tracker• Antenova GPS Chip• Beaglebone

Connections• Purpose

• Track Prometheus• Relay live data to ground

station

Finished Board

Eagle File

Schematic

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Pressure Sensors• Components

• ADC• Op Amps• Pin outs to Beaglebone

cape• Purpose

• Amplify and convert analog pressure data

Eagle FileFinished Board

Schematic

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Dielectrophoresis Components• Components

• Level Shifter• Micro SD• ADXL377 (Accelerometer)• Safety LEDs• Arduino Pro• Camera Connections

• Purpose• The main processing unit

for the dielectrophoresis payload

Eagle File

Schematic

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THANK YOU

QUESTIONS?