PANTHR Hybrid Rocket

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

24. 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

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

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

28. Stability

29. Stability

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

31. Fin Design

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

33. Nose Cone Experiment

34. 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

35. 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

37. 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

38. 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.

40. Questions?