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NASA USLI – FRR

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

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NASA USLI – FRR

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  1. NASA USLI – FRR VU Aerospace ClubRocket-based Studies of Thermoelectric Exhaust Heat Recovery in Aerospace Engines Design of Solid-State Thermoelectric Engine

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

  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. Thermoelectric Engine Design • 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 TEG

  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. Independent Payload Data Acquisition • RDAS Tiny v4.o data boards • Multiple data channels • G-switch • USB connection • Aerocon 1.3H thermocouple boards • Amplify voltage • Linearize Data

  8. Data Acquisition Layout Power Supply RDAS Thermocouple Board Thermocouple Resistor Bank Vout TEG + - + - + - 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. Current Rocket Dimensions & Weight • Length: 124 in • Diameter: 6.3 in • Span Dia.: 26.1 in • CG: 82.9 in • Stability Margin: 2.46 • Launch Mass: 42.9 lb • Thrust-to-Weight: 4.87 • CP: 97.9 in Avionics Bay Electronics Bay Center of Gravity Drogue Parachute Main Parachute Center of Pressure Forward Tube (Recovery Systems) Aft Tube

  11. Rocket Materials • Dyna-wind body tube • Nomex Honeycomb fins laminated in carbon fiber • AV Bay located in coupler tubing

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

  13. Recovery System • Dual deploy recovery system • 12’ main chute with 36” drogue chute • 17 ft/s descent rate • Fireball used to prevent zippering • Black powder (4F) charges • Fore: 4 grams • Aft: 6 grams • Redundant systems using PerfectfliteminiAlt/WD

  14. Approach to Design of Payload • Two primary concerns: • Thrust Loss (if any) • Heat Transfer (necessary) • Addressing concerns • Static testing • Thrust effects of payload • TEG location • Number of TEGs • Preliminary Flights • Safe take-off velocity • Full-scale flight

  15. Payload Requirements • Maximize ΔT while minimizing thrust loss • Minimal Impact on Rocket Stability • Minimal impact on thrust to weight • Eliminate thrust loss due to the Krushnic Effect. • Optimal jet interference with negligible thrust loss

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

  17. Static Testing – Thrust Loss • Motor: Pro 38 I 212-ss • Payload: press fit to motor retainer ring • Blue: Control • Pink: L>D, ¼ inch vents, 30% of anticipated thrust • Yellow: L<D, ½ inch vents, 90% of anticipated thrust

  18. Mitigation of Krushnic Effect • Maximizing ventilation at vena contracta minimizes any thrust loss from Krushnic Effect • Thrust losses seen were likely due to interference, not Krushnic Effect

  19. Static Testing – 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 • Thermocouple to measure payload temperature • Slot ventilation to eliminates Krushnic Effect

  20. Thrust Loss and Heat Transfer • No appreciable loss of thrust • Slot venting negates Krushnic Effect • Pink: Published Thrust Curve • Blue: Thrust measured • ΔT of 30°C (yellow)

  21. Improving Temperature Gradient • Payload Lengthened to increase jet interference • Motor J99 • Payload: 2.5’’ diameter 7’’ long aluminum tube, 3/8’’ diameter 1’’ long slots • Minimal thrust losses seen • Temperature exceeded 600°C • Lengthening payload increases temperature gradient without increasing thrust loss

  22. Full Scale Test FlightFeb 13 – Simulated Payload Successes: • Acceptable rail speed, launch and recovery • Simulated Payload retained well and did not affect flight • Based off of ground based testing Problems: • Fin separation • Premature release of main parachute

  23. Feb 13 Trajectory

  24. Flight Revisions • Revised fin design • Eliminate Aft Sections • Eliminates Flutter • Carbon Fiber fillets to ensure retention • Shear pins to secure recovery systems • RDAS for time synchronization of flight and data streams

  25. Static Testing – Power Generation • Motors: J90 & J99 • Payload: 2.5’’ diameter 8’’ long aluminum tube, ½ ’’ diameter 1.37’’ long slots • TEG and 3 ohm resistor placed for power generation • Cooling fins added for heat sink • VASA fan simulates flight conditions

  26. Power generation results

  27. Flight Trajectory • Tests show no thrust loss • RockSim appears to overshoot based on full scale flight • New trajectory simulator developed to vary payload mass and coefficient of drag

  28. Flight Profile • Lower Bound: Coefficient of drag = 0.8 Payload mass = 2 kg • Upper Bound: Coefficient of drag = 0.6 Payload mass = 1 kg • Payload mass: 1.8 kg (~4 lbs) • Anticipated apogee: 1610 m (5280 ft)

  29. Payload Design – 3/20 • Aluminum • L = 8.5 inches • Minimum wall thickness: .06 inches • Slot venting • 6 TEG, 2 sets of 3 • Heat sink fins • Total weight: 4 lbs

  30. Integration • RDAS boards for data collection • Each board has 3 TEGs and 1 thermocouple • Welded retainer ring for housing • Wiring running through PVC tube for protection

  31. FRR Test FlightMarch 20 • Fin changes from Feb 13 flight • Eliminated flutter by trimming fins • Thermal data collected by RDAS • TEG power generation • Thermocouple readings • Observations on flight: • Payload required extensive cleaning • Slower descent rate desired to protect payload • Use 12 foot main parachute instead of 10 foot Apogee: 5303 ft

  32. USLI Competition Design Changes • TEGs placed on aft row due to higher temperatures measured • Smaller heat sinks • 12 foot main parachute for slower descent rate to protect valuable payload upon landing

  33. Educational Outreach • Dyer Observatory • March 2 • April 6 • Vanderbilt Engineering Open House • April 10

  34. Rocket Safety • Static fires to characterize thrust profiles • 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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