Transcript FLIGHT READINESS REVIEW (FRR) Charger Rocket Works University of Alabama in Huntsville NASA Student Launch 2013-14 Kenneth LeBlanc (Project Lead) Brian Roy (Safety Officer) Chris Spalding (Design Lead) Chad.

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FLIGHT READINESS
REVIEW (FRR)
Charger Rocket Works
University of Alabama in Huntsville
NASA Student Launch 2013-14
Kenneth 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
Payload
Description
Nanolaunch 1200
Record flight data for aerodynamic coefficients
Dielectrophoresis
LHDS
Supersonic Coatings
Use high voltage to move fluid away from container walls
Payloads Here
Detect and transmit live data regarding landing hazards
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
Total Rocket Cost: $3,217
Propulsion
$823.91
Recovery
System
$204.53
Hardware
$1,212.19
Payload
$976.98
$-
$200.00
$400.00
$600.00
$800.00
Cost
$1,000.00
$1,200.00
$1,400.00
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DESIGN
Team Members:
Chris Spalding - Team Lead
Andrew Mills - Prototype Assembler
Jordan Lee - Designer
Josh Thorne - Coatings
David Zaborski - Recovery Designer
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5
1
2
3
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Hardware Changes:
Design Details:
1. Printed Nylon
2. 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
• 34lbs
• 40Gs acceleration
• Geometric similarity to NASA
Nanolaunch prototype
• Nanolaunch team requested
maximum use of SLS printed
aluminum
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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
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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
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Recovery Package
Shock cord, quick links, ring
nut
Tension from ring nuts, aerodynamic forces
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Interfaces (2)
#
Component
Interface
Load Locations
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Payload
Slipped onto payload shaft,
constrained between nuts and
bulkhead
Acceleration forces, passed through payload
shaft
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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
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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
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Motor case
Threaded to payload shaft
Outside manufacture; loaded in designed
manner
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Fins /Fin Brackets
Bolted to lower body tube, Tnuts inside body tube.
Aerodynamic acceleration forces, resulting
tension from body tube.
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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
Currently have sets of fin brackets in abs plastic and fiber reinforced nylon.
• ABS has been proof loaded to 75 lbs
• 3D printed Laser sintered nylon brackets have been
ordered
• Bolted to body
• Binding post fin attachment
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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
Old Design
• 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.
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Pitot Probe
Manufactured out of glass reinforced nylon.
New Design
• 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
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Vehicle Success Criteria
Requirement
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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ANALYSIS
Team Members:
Chad O’Brien - Team Lead
Sarah Sheldon - Design Analyst
Garrett Holmes - LHDS Analyst
Tryston 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 Criteria
Requirement
Success Critera
Verification
Safe Launch Operations
Vehicle maintains safe heading and travel.
Achieve Sustained Transonic Flight
Vehicle Launch does not cause injury to
launch crew or bystanders.
Launch velocity must exceed minimum speed
for stable flight
Stay within Mach 0.7-1.4 and collect usable
data
Stay within 750 ft/s – 1400 ft/s
Supersonic flight reached
Reach above Mach 1
Fin Design Supports Supersonic Flight
Fins should be ready to fly after short post
flight inspection and new flight preparations
without modification.
Lands within 5000 feet of the launch tower
Aerodynamic Stability for launch
Transonic flight data
Meets Drift Requirement
Safe Ground Energy Impact Levels
Recoverable and Reusable
Components must impact with less than 75
feet pound force
Rocket can be launched again without
significant alteration
Visual observation of vehicle behavior off the
rail coupled with empirical data from sensors.
Receive readable data from accelerometers
and pitot pressure sensors
Data from accelerometers or pitot pressure
sensors
Data from accelerometers or pitot pressure
sensors
Visual inspection and simple sturdiness test
to ensure fixtures and material are ok to fly.
Tracker and GPS will be used to verify
position on landing and a distance from
launch sight will be calculated.
Review of flight data to see impact speed.
Rocket can be launched again in same day
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FLIGHT SIMULATIONS
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Mass Statement
Subsystem
Mass (lbs)
Payload
2.93
Recovery
4.7
Airframe
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Motor Case
7.0
Propellant
7.4
Total Dry System
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Total Wet System
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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
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Prometheus’ Static Margin
CG at 85.8”
• Launch Static Margin – 1.7
• Burnout Margin – 4.5
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:
𝜇 = 15,700𝑓𝑒𝑒𝑡
𝜎 = 1700 𝑓𝑒𝑒𝑡
Mach:
𝜇 = 1.39
𝜎 = 0.15
Accleration:
𝜇 = 37.9 𝐺
𝜎 = 4.4 𝐺
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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 directly
affects the radial landing
distance.
• Dependent on high speed
coefficient 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 Fin
Given:
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
𝐹𝑐𝑟𝑖𝑡 = 𝐹𝑚𝑜𝑡𝑜𝑟 = 4770𝑁
𝐺 = 5𝐸5 𝑝𝑠𝑖
𝜈 = 0.1
𝐸 = 1100 𝑘𝑠𝑖
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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 Materials
Part
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
Eye
bolt
L.H.D.S
Tethers
• Stage 3
• Main parachute falls out
deployment bag/burrito
Black Powder Charge
Drogue
Main Parachute
In
Deployment
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
Data Dropout
• 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
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
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
Wind Speed Range (MPH)
Average Drift (feet)
3-4
1388
8-14
3358
15-25
5962
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TESTING AND
VERIFICATION
Brian Roy – Safety Officer
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Testing Procedures
Review of
Procedures by
PRC Staff
Develop
Operating
Procedures
Test
Requirement
Identified
Procedure
Approval by
PRC Director
Identify Red
Team
Members for
Test
Review of
Operating
Procedure with
Red Team
Testing
Approval of
Red Team
Members
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Subscale Testing and Results
Sub-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 Electronics
Ensure 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 Design
Verify 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 Gloading (~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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PAYLOADS
Team Members:
Wesley Cobb - Team Lead
Bronsen Edmonds - Sensor Developer
Tyler Cunningham - Dielectrophoresis
Shawn Betts - LHDS
Samuel Winchester - Tracking
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Payload Integration
LHDS
Aerodynamic
Coefficient
Payload
Dielectrophoresis
and 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
69
Nanolaunch Rocket Rotation Verification
•
•
•
•
Verified using FFT
First peak -> Due to offset
Low frequency = 1.2 Hz
Fast frequency = 11Hz (Low amplitude noise)
70
High G Accelerometer Data(Ground Test)
71
CG Transducer Pressure Data
(Used for Calibration) Setra 830E
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CG Pressure Sensors Calibration Curves
73
Transducer Uncertainty Results
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
• Ways to decrease uncertainty:
• Optimize gain resistor value
• Measure more data points in the Vacuum region to improve calibration curve
• Purchase Higher Precision Transducer
74
Nose Cone Shock Interactions for
Pressure Sensor Consideration
• Bow Shock Due to Blunt Tip
• Measure Stagnation Pressure and Static Pressure After the Shock
75
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
76
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:
77
Nanolaunch Success Criteria
• Objectives: Meet Team/NASA SLI Requirements and
Verify Those Were Met
Requirement
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
78
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
79
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
80
Experimental Changes
• Electrode configuration: from parallel
electrodes, to annular electrodes
• Voltage increase from 7kV to ~12kV
2012-2013
Configuration
2013-2014
Configuration
81
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
82
DEP HV Output Test
This 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
83
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.
84
Supersonic Paints and Coatings
• Urethane
• Excellent retention
• Abrasion resistant
• Smooth Coating
Urethane
Epoxy
• Epoxy Primer
• Low film build
• Excellent adhesion
• Rough Coating
• Thermal tape
• 3-5 second reaction time
• Changes color at specific
temperatures
• Excellent Adhesion
Epoxy
85
SPC Testing
• Oven Testing for Temperature tape
• Calibration of tape
• Temperature sensitivity
• Reaction time
• Flight Test
• Subscale Test Flight
• Full scale test launch
86
Success Criteria of Paints and Coatings
Requirement
Success Criteria
Verification
Even 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
application
High heat resistant
Coating unscathed
from thermal loads
No 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
87
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 LHDS
Radio Shown
88
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 Field
Manchester, TN
After Edge Detection
89
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.
90
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.
91
LHDS Success criteria
Requirement
Success Critera
Verification
Recoverable and Reusable
If the Payload can be removed and replaced
between subsequent flights
The mass simulators or payload does not
hinder the Rocket’s takeoff or landing
If no damage to the rocket is done by the
mass simulators
If the RF antenna can communicate
successfully without impedance from the
design
If the camera can take pictures of the ground
without any obstruction from the Rocket
body after deployment
RF module communicates with GPS module
and ground
If electronics are functional and record data
properly
There are no missing screws or bolts or tools
necessary to fix the LHDS.
If the payload remains inside its design, and
the Rocket launch and landing are successful.
If any scratches or dents are visible inside or
outside the rocket near the payload
If a signal is reached from the ground base.
Sustainable
Non-damaging
Communicable
Camera Visibility
Communicable GPS
Functioning Electronics
If the pictures have desired amount of
ground-landing in them to be able to verify
landing hazards
If GPS location is communicated to ground
If data is recorded and readable for analysis
92
Beaglebone Cape
• Components
• ADLX 377 (Analog
Accelerometer)
• L3GD20 (Gyroscope)
• ADLX 345 (Digital
Accelerometer)
• IC^2 connection to other
boards
Final Product printed by OHS Park
• Purpose
• Detect launch and
initiate Data Acquisition
• Take measurements
CAD Drawing
93
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
94
Pressure Sensors
• Components
• ADC
• Op Amps
• Pin outs to Beaglebone
cape
Finished Board
Eagle File
• Purpose
• Amplify and convert
analog pressure data
Schematic
95
Dielectrophoresis Components
• Components
• Level Shifter
• Micro SD
• ADXL377 (Accelerometer)
• Safety LEDs
• Arduino Pro
• Camera Connections
Schematic
• Purpose
• The main processing unit
for the dielectrophoresis
payload
Eagle File
96