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NASA USLI – CDR

NASA USLI – CDR VU Aerospace Club Rocket-based Studies of Thermoelectric Exhaust Heat Recovery in Aerospace Engines. Project Overview . Recover waste heat from exhaust Use thermoelectric generators (TEG) Initial Testing for: Recovered Heat Thrust Losses

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NASA USLI – CDR

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  1. NASA USLI – CDR VU Aerospace ClubRocket-based Studies of Thermoelectric Exhaust Heat Recovery in Aerospace Engines

  2. Project Overview • Recover waste heat from exhaust • Use thermoelectric generators (TEG) • Initial Testing for: • Recovered Heat • Thrust Losses • Use rocket to simulate flight conditions of jet

  3. Application • Use thermoelectric generators to recover waste heat coming off of jet engines • Commercial jets run for hours at a time • Can recover energy to power on-board systems during flight • Small improvements = big benefits on large scale

  4. Motivation • Energy efficiency is important global topic • Typical jet engines run at 35% efficiency • Small increases in efficiency can pay large dividends • Test for net increase in efficiency • Rocket is cheapest means of testing theory while achieving in-flight operating conditions

  5. Payload Description • Thermoelectric Generators (TEGs) will recover waste heat in exhaust flow • Attach TEG assembly on aft end of rocket • Voltages read and stored on-board by data logger

  6. Thermoelectric Generator • Allows for direct conversion between thermal and electrical energy • Function through motion of excess electrons and the Seebeck Effect • Very reliable since no moving parts • Lightweight

  7. Data Acquisition • SparkFunLogomatic V2 microSDDataloggers on board • Each board has up to 10 channels • Low cost and recording to SD makes accessing data simple • Aerocon Thermocouple Sensor Boards

  8. Data Acquisition Layout SD Cardash MCU Vout A/D Resistor bank TEG + - + - TEG

  9. Vehicle and Motor Requirements • 6 inch body tube diameter • Sufficient power to ensure safe thrust to weight • Sufficient stability to account for additional payload

  10. Scaled Model - Prototype • 1:6 scale model of final rocket design • Created using ProE and rapid prototyping machine • Vanderbilt Wind Tunnel for aerodynamic testing

  11. Rocket Dimensions & Weight • Length: 125 in • Diameter: 6.3 in • Span Dia.: 26.3 in • CG: 101.3 in • Stability Margin: 3.09 • Launch Mass: 40.0 lb • Thrust-to-Weight: 5.23 • CP: 82.4 in

  12. Rocket Materials • Dyna-wind body tube • Carbon Fiber Composite Fins for added stability • AV Bay located in coupler tubing

  13. Motor • Cesaroni Pro98 L610 • Long burn (8.1s) • Aeropack Retainer • Thrust to weight: • 5.23 (initial) • 3.7 (average)

  14. Recovery System • Dual deploy recovery system • 10’ main chute with 24-36” drogue chute • 22 ft/s descent rate • Fireball used to prevent zippering • Black powder (4F) charges • Fore: 4 grams • Aft: 6 grams • Redundant systems using PerfectfliteminiAlt/WD

  15. Motor Testing • Motor • Static testing • Thrust effects of payload • TEG location • Number of TEGs • Preliminary Flights • Safe take-off velocity • Full-scale flight

  16. Thrust Loss • Loss due to overexpansion: Krushnic Effect • Loss due to exhaust jet interference • Interference is required for heat transfer

  17. Payload Requirements • Maximize ΔT while minimizing thrust loss • Minimal stability loss • Minimal impact on thrust to weight • Minimal thrust loss due to the Krushnic Effect.

  18. Payload Testing • Estimation of effects on motor thrust and payload heat • G-79 test motor • Run 1: Estimation of loss of thrust • Run 2: ½ inch ventilation holes added

  19. Results • No appreciable loss of thrust • Krushnic Effect not encountered due to large diameter of payload • No heat transfer due to large diameter of payload

  20. Krushnic Effect • Motor: Pro 38 I 212ss • Payload: fit to motor nozzle diameter • Run 3: Control (above) • Run 4: L>D, ¼ inch vents • Run 4 has 30% of anticipated thrust

  21. Payload Restrictions • Run 5: L<D, ½ inch vents • Run 5 has 90% of anticipated thrust • Thrust loss likely due to interference, not Krushnic Effect • L=D maximum, maximize ventilation

  22. Minimizing Thrust Loss Maximizing Heat Transfer • Motor: J99 • Payload: 2.5’’ diameter 5’ long aluminum tube, 3/8’’ diameter 1’’ long slots • Long duration burn to model rocket flight • Slot ventilation to minimize Krushnic Effect

  23. Thrust Loss vs. Heat Transfer • No appreciable loss of thrust • Slot venting negates Krushnic Effect • Maximum ΔT of 30°C

  24. J99 Motor Results

  25. Thrust Loss and Flight Trajectory • Tests show 10% loss in thrust possible due to interference • Thrust loss alters flight trajectory • Trajectory estimation using RockSim • Without Thrust Loss: • Thrust to Weight: 5.23 • Apogee:~6400 feet • With 10% Thrust Loss: • Thrust to Weight: 4.7 • Apogee:~5400 feet

  26. Flight Profiles Without Thrust Loss With 10% Thrust Loss

  27. Future Tests • Optimize heat transfer • Add TEG to ground testing to estimate power output • Add fins on payload to improve temperature difference

  28. Payload Design • Aluminum • L=6 inches • Multiple vents • Hexagonal shape • Currently 1.65 lbs (including retaining ring mounting)

  29. Integration • Utilizing microSD to log data on-board • System small enough to make redundant collection viable • Welded retainer ring for housing • Mounting bolts to retain aft payload

  30. Anticipated Test Flight Dates • Elizabethtown, KY- Feb 13th • Memphis, TN or Birmingham, AL - Feb 20th • Manchester, TN - March 13th (Pending FAA Waiver)

  31. Educational Outreach • Dyer Observatory • Nashville Area Primary Schools • Vanderbilt Engineering Open House

  32. Rocket Safety • Static fires to characterize thrust loss • All components used to industry standards • Stability margin within safe range • Scaled and preliminary flights • All codes and laws followed during all team events

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