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PANTHR Hybrid Rocket

PANTHR Hybrid Rocket

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PANTHR Hybrid Rocket

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  1. PANTHRHybrid Rocket Final Design ReviewDecember 6th 2006

  2. PANTHR Team Members Glen Guzik Niroshen Divitotawela Michael Harris Bruce Helming David Moschetti Danielle Pepe Jacob Teufert

  3. Current Division of Labor • Hybrid Motor Design- Niroshen Divitotawela- Michael Harris- Jacob Teufert • Aerodynamics and Flight Stability- Bruce Helming- Danielle Pepe • Payload and Recovery- Glen Guzik- Bruce Helming - Danielle Pepe - Michael Harris • Structural Analysis- David Moschetti - Niroshen Divitotawela • Safety and Logistics-David Moschetti-Glen Guzik

  4. Primary Project Objectives • Build hybrid rocket motor - paraffin fuel (CnHm; n~25, m~50) - nitrous oxide oxidizer (N2O) • Conduct static test fire • Complete fabrication of rocket • Launch rocket to an altitude of ~12,000 ft. • Collect various in-flight data- acceleration curve- flight trajectory- altitude at apogee- onboard flight video

  5. The Paraffin Advantage Advantages of Paraffin • High Regression Rate • Practical Single-Port Design • High Energy Density (~same as kerosene) • Inexpensive • Non-toxic Advantages of Nitrous Oxide • Available • Inexpensive • Self-Pressurizing

  6. MOTOR EXPLODED VIEW OXIDIZER TANK ABLATIVE LINER INJECTOR FUEL GRAIN NOZZLE COMBUSTION CHAMBER

  7. Oxidizer Fill and Ignition System • Fill internal oxidizer tank via external, commercial nitrous-oxide tank. • Light solid propellant ignition charge via electric match.

  8. Trajectory Analysis • 1 Degree of Freedom • Explicit First-Order Finite Difference Method • Thrust and Mass=f(t) • Drag=f(v) • Density=f(h)

  9. Regression Rate • Use regression rate formula for hybrids • a = .155, n=.5 [1] • Regression Rate = 1.98 mm/s [1] AA283 Aircraft and Rocket Propulsion – Hybrid Rockets. Stanford University Department of Aeronautics and Astronautics. 2004

  10. Combustion Chamber Dimensioning From Trajectory Analysis: • Average Mass Flow: 0.375 kg/s • Burn Time: 4 s From Literature Review: • Regression rate as f(dm/dt) • Oxidizer/Fuel Ratio Results: • Grain Thickness (d=rtb) • Grain Length

  11. Combustion Chamber Dimensions • Grain Length: 4.2” • Grain Thickness: 0.68” • Chamber Wall Thickness: 1/8” • Ablative Liner Thickness: 1/8” 3.0” Combustion Chamber Ablative Liner Fuel Grain 4.2” 1.14”

  12. Combustion Chamber Thermodynamic Properties From Analysis • Adiabatic Flame Temperature: 3800K From Literature Review • Paraffin Flame Temperature: 1700K For Design • Average Value: 2750K

  13. Nozzle Design • Method • Decided to expand the flow to sea-level pressure. • Use of isentropic relations • Find the Area Ratio • From trajectory computation make use of estimate of mass flow rate.

  14. Non-Ideal Expansion

  15. Specifications • Conical Nozzle • Ae/A* = 3.64 • Divergence Angle of 8o • Length 3.87” • Weight 1.13 lbs.

  16. Trade Study • Several Alloys were compared in the decision process for the material of the tubing needed for the tank. • Al 6061-T6 was observed to be the best metal to use considering cost and strength. Ratings were acquired by the Hadco Aluminum website.

  17. Structural Analysis • Most severely stressed components are the Combustion Chamber and Oxidizer Tank • Wall Thickness was calculated using hoop stress equation With F.S. of 2:

  18. Max hoop stress (ANSYS) = 9640 psi • Max hoop stress (Theory) = 10000 psi

  19. Structural Analysis • We are using 8 bolts the attach each bulkhead • Each bolt is made of 1022 Carbon Steel • The Allowable Shear for each bolt is 29,000 psi Total Force acting on Bulkheads: Shear on each Bolt:

  20. Structural Analysis Bearing Stress Yield: • The Bearing Stress was calculated for the aluminum tube using the force of the load distributed to each bolt • With that the calculation divides the load by the thickness of the wall, diameter of each hole, and the number of bolts • The allowable was found to be 1.5 times the allowable Tensile strength Total Bearing Stress:

  21. PayloadLayout

  22. Payload Data Collection • Acceleration versus time in 3 dimensions • Pressure versus time • Flight video at 30 FPS 352 x 240

  23. Payload Drop Test • Launch zone, “The Grid” • Impact velocity of up to 25 ft/s • Equivalent to a drop from 10 ft • Survive landing on: trees, rocks, grass, and asphalt

  24. Stability • Maintain the Static Margin • Options: - Under-damped - Neutral - Over-damped • Current Configuration: - over-damped http://www.rockets4schools.org/education/Rocket_Stability.pdf

  25. Stability • Subsonic flight allows use of Barrowman Method • Xcp (Tail reference) = 14.45 inch • Xcg (Tail reference) varies between 34.26 – 38.25 inches

  26. Stability

  27. Stability

  28. Fin Design • 3 different fin designs based on initial rocket plans • Flutter conditions accounted for • Wind tunnel testing was performed

  29. Fin Design

  30. Fin Specifications • Dimensions based on flutter analysis, testing, and stability calculations: • Cr = 6” • Ct = 2.5” • S = 4” • t = 0.167” Ct S Cr

  31. Nose Cone Experiment

  32. Types of Nose Cones 1) Elliptical 2) Conical • They both have low drag characteristics in low-transonic Mach regions. • Elliptical Shape • Total Drag From Experiment = 0.029 • Small Length and Weight decrease Static Margin • Conical Shape • Total Drag From Experiment =0.041 • Length and Weight increase Static Margin Final Choice: Elliptical http://myweb.cableone.net/cjcrowell/NCEQN2.DOC

  33. Recovery Nosecone • Barometric Altimeter • Drogue Chute – Deploys at apogee • Main Chute – Deploys when altimeter detects specified altitude (~1500ft) Drogue Parachute Main Parachute Cut Away View

  34. Spring Semester 2007 Milestones • February 12th: Complete Motor Construction • February 18th: Static Test Fire • February 26th: Complete Payload Construction • March 13th: Payload Drop Test • March 22nd: Rocket Fabrication Finalized • Launch 2nd Week of April

  35. Safety Plan Main Risks • High Pressure Systems • Chemicals/Flammables • Test Fire and Launch Procedures • Construction Mitigation Plan • Currently working with the University Safety Office on developing procedures for handling, construction, and launch of the rocket.

  36. Questions?