Hybrid Rocket Final Presentation
- 1. June 7, 2016 Final Presentation
- 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. 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. 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. 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. Hot End 1 314-1 Max Flansberg Anthony Harteloo David Ha
- 7. Subsystem Layout Post Combustion Chamber Fuel Pre Combustion Chamber Injector
- 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. 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 – Highly reliable – Increase specific impulse Source: https://www.youtube.com/watch?v=Bgy C1jXTY4c
- 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. Results • Fuel – Specific impulse 210 s • Igniter – Chamber temperature > 823K – Burn time > 2s • Injector – 54% increase in thrust – 32% increase in specific impulse – Substantial improvement to combustion efficiency – Subsystem weight within tolerance
- 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. Hot End 2 314-4 Ben Smucker Frank Huynh Kyle Fox
- 14. Overview • Injector Integration (Ben) • Combustion Chamber (Frank) • Nozzle (Kyle) • Relevant Customer Requirements – Lightweight – Durability of parts – Conform to size restraints
- 15. Injector Integration • Manifold /Combustion Chamber – 4-40 socket head cap screws – Loaded in tension • Injector Attachment – Lip – No fasteners in injector
- 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. 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. 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. Cold End 314-3 Joshua Laas Luis Mendoza Nigel Swehla
- 20. Overview • Driving CRs – Safe pressurization – Non-corrosive – High Isp • Proposed problems – Safety – Spatial constraints – Acceptable pressure & mass flow rate • Subsystem solutions – Oxidizer/pressurant – Feed system components – Remote oxidizer priming
- 21. Oxidizer/Pressurant • Nitrous Oxide – Non-cryogenic – Common – Non-toxic – Easy storage – Better performance – Cooling • Self-pressurizing – Single tank – No pressurant – Acceptable pressures
- 22. Feed System Components • Single tank configuration – Light weight – Smaller size • Material selection – Stainless steel – Brass – Aluminum • Physical properties – Length: 44.90 in – Width: 4.65 in – Weight: 12.30 lbs (dry)
- 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. Results Passed: 10 out of 12 testing procedures Failed: Specific Impulse Performance Test Target: 221-270 s Achieved: 210 s Reason: Ambient temperature of 56℉ during test, optimal oxidizer temperature between 70 and 74℉ Oxidizer Pressure Test Target: 760-1500 psi Achieved: 550 psi Reason: Ambient temperature of 56℉ during test, optimal oxidizer temperature between 70 and 74℉
- 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. Vehicle Engineering 314-2 Rodney Fischer Krissy Kellogg Parker Weide
- 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. 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. Aerodynamics Fins • Trapezoidal with Airfoil • T6 6061 Aluminum • Manufactured on CNC 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. 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. 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. 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. Launch Rail 322 Moises Higgins Michael Arthur Loren Valiente Nathan Leendertse Sam Monroe
- 34. Launch Rail: Project Description
- 35. Launch Rail: Design Development
- 36. Launch Rail: Design Solution
- 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 $321.19 (approx. 394.71) Avionics $158.99 Total (goal $5000) $ 5046.61 Launch Rail Total (goal $2000) $1990
- 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 collaborationRecommendations: • Divide up responsibilities differently • More flight-like testing • Larger motor
- 41. Conclusion Sincere thanks to OSU AIAA, John Lyngdahl, Steve Cutonilli and Dr. Squires for their support!
- 42. Questions?