1 FLIGHT READINESS REVIEW FRR Charger Rocket Works

1 FLIGHT READINESS REVIEW (FRR) Charger Rocket Works University of Alabama in Huntsville NASA Student Launch 2013 -14 Kenneth Le. Blanc (Project Lead) Brian Roy (Safety Officer) Chris Spalding (Design Lead) Chad O’Brien (Analysis Lead) Wesley Cobb (Payload Lead)

2 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

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

4 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

5 On Pad Cost Total Rocket Cost: $3, 217 Propulsion $ 823. 91 $ 204. 53 System Recovery 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

6 DESIGN Team Members: Chris Spalding - Team Lead Andrew Mills - Prototype Assembler Jordan Lee - Designer Josh Thorne - Coatings David Zaborski - Recovery Designer

7 1 2 3 4 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 • 34 lbs 5 • 40 Gs acceleration • Geometric similarity to NASA Nanolaunch prototype • Nanolaunch team requested maximum use of SLS printed aluminum

8 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

9 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, Tnuts 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

10 Thrust Ring • Machined 5086 Aluminum • Will be Analyzed with FEA

11 Fin Assemblies Currently have sets of fin brackets in abs plastic and fiber reinforced nylon. • ABS has been proof loaded to 75 lbs • 3 D printed Laser sintered nylon brackets have been ordered • Bolted to body • Binding post fin attachment

12 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

13 Payload Shaft • 7075 -T 6 Aluminum threaded shaft 3/8 -16 • Preloaded in tension • FEA and hand calculations show significantly over strength requirements

14 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

15 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

16 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

17 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

18 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 3 D printing techniques • Helps fulfill our Nanolaunch request to explore selective laser sintering in original ways.

19 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 3 D printing

20 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 in Practice procedures to get team and ready to fly in less than 3 hours from arrival at fluent in the assembly reasonable time launch field 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

21 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

22 GENERAL ROCKET MISSION PERFORMANCE CRITERIA

23 General Rocket Flight Performance Criteria Requirement Success Critera Verification Safe Launch Operations 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 Reach above Mach 1 Vehicle maintains safe heading and travel. Aerodynamic Stability for launch Transonic flight data Achieve Sustained Transonic Flight Supersonic flight reached Fin Design Supports Supersonic Flight Meets Drift Requirement Safe Ground Energy Impact Levels Recoverable and Reusable 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 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 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

24 FLIGHT SIMULATIONS

25 Mass Statement Subsystem 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

26 Prometheus Simulation • Rock. Sim Software Package • Motors • Primary: CTI 4770 – 98 mm • Secondary: Aero. Tech K 1499 – 75 mm • Estimated Dry Mass at 27 lbs • Launch Conditions for Salt Lake City • ASL – 4210 feet • Temperature – 72 ˚F

27 Final Motor Selection - CTI M 4770 -P • • ISP – 208. 3 s Loaded Weight: 14. 4 lb Propellant Weight: 7. 3 lb Max Thrust: 1362 lbf

28 Prometheus’ Static Margin CG at 85. 8” • Launch Static Margin – 1. 7 • Burnout Margin – 4. 5 CP at 93. 6”

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

30 Prometheus’ Static Margin • Pre-Launch Static Margin: 1. 7 • Burnout Static Margin: 4. 5

31 Monte Carlo Analysis

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

33 Variation in Flight Time • Time variance directly affects the radial landing distance. • Dependent on high speed coefficient of drag for drogue

34 Plan B Motor: Aerotech K 1499 • Altitude – 2100 feet • Velocity – Mach 0. 25 • Acceleration – 16 G’s

35 FIN FLUTTER ANALYSIS

36 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

37 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 = b 2/S • l = Ct/Cr •

38 Prometheus Fin Given: Variable Value Units Cr 8. 31 in. Ct 4. 75 in. t 0. 17 in. b 4 in. G* 5. 00 E 5 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

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

40 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

41 Buckling Analysis •

42 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

43 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

44 Main Parachute • • • 12 Feet Semi-Hemispherical Ripstop Nylon Custom Seam 14 Gores Shroud Lines: 0. 125 in x 550 lb Paracord

45 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

46 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

47 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 Tether s Black Powder Charge Drogue Main Parachute In Deployment bag/Burrito

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

49 GPS Tracking • GPS Module: Antenova M 10382 -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 m. A current draw • Since CDR, redundant GPS Unit: “Tagg Pet Tracker” no longer included

50 Radio Transmission • RF Module: XBee-PRO XSC S 3 B • 900 MHz transmit frequency • 20 Kbps data rate • 9 mile Lo. S range • 250 m. W transmit power • 3. 3 VDC supply voltage • 215 m. A current draw • 1. 5+ hr battery life at max sensor sample rate • Laptop ground station

51 GPS/RF Module Ground Testing Data Dropout • Stationary ground station • Transmitter driven away from receiver, increasing Lo. S obstructions • Obstructions were increased until data dropout • Test was a success in worst case scenario terrain conditions Lo. S Ground Station

52 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 M 10382 -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

53 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

54 TESTING AND VERIFICATION Brian Roy – Safety Officer

55 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

56 Subscale Testing and Results Sub-Scale Flight Test Matrix Type of 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) Test hardware that will allow for a Recovery System single separation dual deploy Successful (4/12/14) Hardware setup in full-scale vehicle. Parachute Design Verify construction techniques are adequate and determine effective drag coefficient. High Acceleration Ensure that avionics will survive Flight (40+ G’s) launch forces of full-scale. Successful (2/22/14) Successful (3/8/14)

57 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.

58 Subscale Flight #1 Data • • Apogee: 1, 573 feet AGL. Max Velocity: 279 ft/s. Time of Flight: 63. 9 seconds. Recorded Using a Perfect. Flite SL 100

59 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-600 R. • Flight Results: Main parachute did not deploy. Nanolaunch payload prematurely triggered, data not usable.

60 Subscale Flight #2 Data • • Apogee: 4, 156 feet AGL. Max Velocity: 597 ft/s. Time of Flight: 128. 6 seconds. Recorded Using a Perfect. Flite SL 100

61 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).

62 Subscale Flight #3 Data • • Apogee: 7, 758 feet AGL. Max Velocity: 1, 208 ft/s. Time of Flight: 210 seconds. Recorded using a Perfect. Flite SL 100

63 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.

64 Prototype Flight #1 Data • • Apogee: 1, 259 feet AGL. Max Velocity: 279 ft/s. Time of Flight: 65. 5 seconds. Recorded using a Perfect. Flite SL 100

65 PAYLOADS Team Members: Wesley Cobb - Team Lead Bronsen Edmonds - Sensor Developer Tyler Cunningham - Dielectrophoresis Shawn Betts - LHDS Samuel Winchester - Tracking

66 Payload Integration LHDS Aerodynamic Coefficient Payload Dielectrophoresis and Aerodynamic Coefficient Payload

67 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

68 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 = 11 Hz (Low amplitude noise)

70 High G Accelerometer Data(Ground Test)

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

72 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 • Po 2/P 1 ratio can be directly looked up to find M 1 and M 2 Atmospheric Pressure Measured Data Calculated Ratios Before Shock After Shock P 1 [Pa] Po 2/P 1 P 2/P 1 M 2 101325 101353. 4 101325 1. 000 0. 020 101325 102036 101325 1. 007 1. 000 0. 100 101325 104190. 6 101325 1. 028 1. 000 0. 200 101325 107853. 4 101325 1. 064 1. 000 0. 300 101325 113134. 6 101325 1. 117 1. 000 0. 400 101325 120193 101325 1. 186 1. 000 0. 500 101325 129240. 4 101325 1. 276 1. 000 0. 600 101325 140548 101325 1. 387 1. 000 0. 700 101325 154453. 8 101325 1. 524 1. 000 0. 800 101325 171371. 3 101325 1. 691 1. 000 0. 900 101325 191801 101325 1. 893 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 Po 2 and P 2 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 Velocity Verification Success Criteria Verification Measure Pitot static Recover pressure data pressure at the nose to from the Pitot static calculate Mach 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 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 7 k. V to ~12 k. V 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 60 k. V

83 DEP Success Criteria Requirement Microgravity environment Success Criteria Verification Reach apogee of flight Retrieve to experience accelerometer data to microgravity determine duration of environment 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 • Epoxy Primer • Low film build • Excellent adhesion • Rough Coating • Thermal tape • 3 -5 second reaction time • Changes color at specific temperatures • Excellent Adhesion Epoxy Urethane 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) • L 3 GD 20 (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

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 Eagle File • Purpose • Amplify and convert analog pressure data Schematic Finished Board

95 Dielectrophoresis Components • Level Shifter • Micro SD • ADXL 377 (Accelerometer) • Safety LEDs • Arduino Pro • Camera Connections Schematic • Purpose • The main processing unit for the dielectrophoresis payload Eagle File

96