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HYBRIDROCKET

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HYBRIDROCKET

  1. 1. June 7, 2016 Final Presentation
  2. 2. Overview Project description Design process Sub-team presentations: • Hot End 1 • Hot End 2 • Cold End • Vehicle Engineering • Avionics • Launch Rail Budget Overall results and recommendations Conclusion, Q&A
  3. 3. Project Description What is a hybrid rocket? spacesafetymagazine.com Project background and purpose: • Sponsored by OSU chapter of AIAA • Design and build a flight-viable hybrid rocket • Research and optimize for the future • Collaborative design experience
  4. 4. Multidisciplinary Collaboration Fuel and Energetics - Chemical Engineers • Fuel composition optimization • Combustion reaction energetics Data Acquisition - Electrical Engineers • DAQ system for propulsion test stand bbc.co.uk Hybrid Rocket Design - Mechanical Engineers ● Hot End 1 - injector, fuel, igniter ● Hot End 2 - combustion chamber, injector manifold, nozzle ● Cold End - oxidizer feed system, remote priming system ● Vehicle Engineering - recovery system, aerodynamics, structures, integration Launch Rail - Mechanical Engineers ● Large enough for hybrid rocket ● Collapsible for travel
  5. 5. General Design Process Team Customer Requirements: 1. Safe 4. Can be assembled quickly 2. Reliable 5. Flight viable 3. Cost-effective 6. Lightweight House of Quality: CRs and ESs Simulation: Solidworks ANSYS OpenRocket NASA CEA Sub-system testing: Sub-team TPs Full-scale testing: Test fires Dry Run Assemblies
  6. 6. Hot End 1 314-1 Max Flansberg Anthony Harteloo David Ha
  7. 7. Subsystem Layout Post Combustion Chamber Fuel Pre Combustion Chamber Injector
  8. 8. Fuel Composition • Function – Provide the chemical energy to propel the rocket • Rocket Fuel Composition – Composed of paraffin wax, corn starch, and aluminum powder – Corn starch plasticizes the paraffin wax – Aluminum powder increases the combustion temperature • Customer Requirements – Thrust to weight ratio – Operating Temperatures – Specific impulse of motor
  9. 9. Igniter • Function – Preheat the combustion chamber – Decompose oxidizer • Powderless solid igniter grain – 65% potassium nitrate, 25% sugar, 10% corn syrup by mass – Ignited using commercial E-match – Located in pre-combustion chamber • Customer Requirements – Suitable chamber temperature – High reliability – Increase specific impulse Source: https://www.youtube. com/watch?v=BgyC1jXTY4c
  10. 10. Injector • Function – Inject oxidizer into combustion chamber – Strong effect on motor performance • Dictates flow field in rocket • Swirling Injector – Increases burning rate of fuel – Increased combustion efficiency over design alternative • Customer requirements satisfied – Increase specific impulse – Combustion efficiency – Reliable – Flight Viable
  11. 11. Results • Fuel – Specific impulse 210 s • Igniter – Chamber temperature > 823K – Burn time > 2s • Injector – 54% increase in thrust – 76% increase in specific impulse – Substantial improvement to combustion efficiency – Subsystem weight within tolerance
  12. 12. Recommendations • Increase melting temperature of the fuel • Scale up motor for competition • Lengthen the post combustion chamber • Shorten the precombustion chamber • Continue researching high energy fuels • Refine igniter grain manufacturing process
  13. 13. Hot End 2 314-4 Ben Smucker Frank Huynh Kyle Fox
  14. 14. Overview • Injector Integration (Ben) • Combustion Chamber (Frank) • Nozzle (Kyle) • Relevant Customer Requirements – Lightweight – Durability of parts – Conform to size restraints
  15. 15. Injector Integration • Manifold /Combustion Chamber – 4-40 socket head cap screws – Loaded in tension • Injector Attachment – Lip – No fasteners in injector
  16. 16. Combustion Chamber • Aluminum Chamber • Initial design failed during testing • Wall thickness: increase from 3.6 mm to 8.0 mm • Maximum temperature: 600F • Tensile Yield Strength: 4640 psi • Design Pressure of 400 psi Initial Design Final Design
  17. 17. Nozzle • Accelerate flow of combustion products - Conical nozzle for simplicity • Rocket may not leave rail safely - Optimized to help meet speed requirement off launch rail • Nozzle Design has uncertainty – Combustion gases require mixing – Design does not change much
  18. 18. Results and Recommendations Results • Maintained integrity during testing • Nozzle produced phenomena indicative of flow acceleration • Assembly fits in the rocket • Some tests failed in order to fit the motor in the rocket. Recommendations • Design for manufacturing (Avoid boring!) • More heat transfer analysis in the combustion chamber walls • Further optimize nozzle for launch altitude
  19. 19. Cold End 314-3 Joshua Laas Luis Mendoza Nigel Swehla
  20. 20. Overview • Driving CRs – Safe pressurization – Non-corrosive – High Isp • Proposed problems – Safety – Spatial constraints – Acceptable pressure & mass flow rate • Subsystem solutions – Feed system components – Oxidizer/pressurant – Remote oxidizer priming
  21. 21. Oxidizer/Pressurant • Nitrous Oxide – Non-cryogenic – Common – Non-toxic – Easy storage – Better performance – Cooling • Self-pressurizing – Single tank – No pressurant – Acceptable pressures
  22. 22. Feed System Components • Single tank configuration – Light weight – Smaller size • Material selection – Stainless steel – Zinc plated steel – Aluminum • Physical properties – Length: 44.9 in – Width: 4.65 in – Weight: • 11.55 lbs (dry) • 19.00 lbs (wet)
  23. 23. Remote Oxidizer Priming • Collapsible arm – Quick disconnect fittings – Worm drive and gear motor – Servo to Initiate • Accessibility – Side of rocket – Door • Material selection – Stainless steel – Electronic disconnect
  24. 24. Results Passed: 10 out of 12 testing procedures Failed: Specific Impulse Performance Test Target: 221-270 s Achieved: 211 s Reason: Ambient temperature of 56℉ during test, optimal oxidizer temperature between 70 and 74℉ Oxidizer Pressure Test Target: 760-1500 psi Reason unknown due to pressure transducer failure. May be estimated with calculations from injector pressure data.
  25. 25. Recommendations For better results: • Test rocket motor vertically • Use double valve carbon fiber tank • Pressurize with inert gas • Decrease plumbing pressure drop – Test larger check valve • Temperature control storage tank For ease of data analysis: • Ensure all sensors work for every test • Advance method for cleaning up data • Create a venturi flow meter https://www.youtube.com/watch?v=ZyfvJF529no For safety: • Keep impressing other hybrid teams by enhancing existing safety mechanisms, preventing spontaneous nitrous oxide reactions http://www.simmonsmfg.com/wp- content/uploads/2012/11/PAGE-5-CC1b. jpg
  26. 26. Vehicle Engineering 314-2 Rodney Fischer Krissy Kellogg Parker Weide
  27. 27. Structures •Rolled carbon fiber body tubes (ICE) •Fiberglass couplers •Aircraft plywood bulkheads, centering rings –Steel dowel pins where necessary •Blue tube motor tube Relevant Customer Requirements: • Integrates with motor assembly, recovery, avionics • Stable throughout all stages of flight • Easy to manufacture
  28. 28. Recovery Parachute: • Modified cruciform, Nylon webbing shroud lines, Kevlar shock cord, stainless steel hardware • Total weight 0.362 kg • Provides 43.4 lb drag force, for a landing velocity of 7.3 m/s Deployment: • Piston • Aluminum charge holder • 1.2 g Triple 7 ejection charge, e-match ignitionRelevant Customer Requirements: • Recoverable with minimal damage • Easy to manufacture • Lightweight
  29. 29. Aerodynamics Fins • Trapezoidal with Airfoil • T6 6061 Aluminum • Manufactured using CNC mill and hand finished Relevant Customer Requirements: • Stable throughout all stages of flight • Recoverable with minimal damage Nose Cone • Von Karman with 5:1 Fineness Ratio • Fiberglass with Aluminum tip • Male mold manufacturing process Relevant Customer Requirements: • Stable throughout all stages of flight • Lightweight
  30. 30. Integration Motor assembly → Plumbing → Tank+Avionics → Recovery → Nosecone Relevant Customer Requirements: • Integrates with motor assembly, recovery, avionics • Requires reasonable assembly time • Easy to manufacture
  31. 31. Results and Recommendations Results • Solutions provided good balance of simplicity and reliable functionality • Passed 10 of 14 testing procedures – True failure: stability margin, pin shear force – Failed body tube bending, body tube buckling due to underestimation of performance (rocket is stronger than anticipated) Recommendations: • Recovery bulkhead e-match connectors • RF transparency to simplify wireless communication • Female mold for nosecone
  32. 32. Avionics Beaglebone Black embedded computer to control ignition, oxidizer valve Redundant Stratologger systems to deploy parachute Xbee wireless module for remote control and communication
  33. 33. Launch Rail 322 Moises Higgins Michael Arthur Loren Valiente Nathan Leendertse Sam Monroe
  34. 34. Launch Rail: Project Description
  35. 35. Launch Rail: Design Development
  36. 36. Launch Rail: Design Solution
  37. 37. Launch Rail: Results Budget
  38. 38. Launch Rail: Recommendations
  39. 39. Budget Rocket Sub-Team Amount Hot End 1 $1406.46 Hot End 2 $1020.95 Cold End $1824.07 Vehicle Engineering $336.14 Data Acquisition $300 (approximate) Avionics $158.99 Total (goal $5000) $ 5046.61 Launch Rail Total (goal $2000) $1990
  40. 40. Overall Results and Recommendations Results: • Flight viable rocket • Launch scrubbed due to time constraints – Avionics communication issue – Ball valve battery issue • Excellent teamwork and collaboration Recommendations: • Divide up responsibilities differently • More flight-like testing • Larger motor
  41. 41. Conclusion Sincere thanks to OSU AIAA, John Lyngdahl, Steve Cutonilli and Dr. Squires for their support!
  42. 42. Questions?

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