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PANTHR Hybrid Rocket Final Design Review December 6 th 2006

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

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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

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Primary Project Objectives Build hybrid rocket motor - paraffin fuel (C n H m ; n~25, m~50) - nitrous oxide oxidizer (N 2 O) 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

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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

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

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

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

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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

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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=rt b ) Grain Length

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

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

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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.

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Non-Ideal Expansion

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

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Trade Study Scale Far Below Average 0 Below Average 1 Average 2 Above Average 3 Far Above Average 4 Material Strength/ Density RatioWeldabilityMachinability Corrosion ResistanceAvailabilityCostScore Aluminum 7075 - T641321213 Aluminum 2024 - T33.52313113.5 Aluminum 6061 - T632234418 Aluminum 6061 - O14131111 Aluminum 6061 - T42.53231112.5 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.

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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:

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Max hoop stress (ANSYS) = 9640 psi Max hoop stress (Theory) = 10000 psi

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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 Shear on each Bolt: Total Force acting on Bulkheads:

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Structural Analysis 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 Bearing Stress Yield: Total Bearing Stress:

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Payload Layout

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

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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

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

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Stability Subsonic flight allows use of Barrowman Method Xcp (Tail reference) = 14.45 inch Xcg (Tail reference) varies between 34.26 – 38.25 inches Center of Pressure X-bar (in)p(x)X*p(x) Nose Cone 326 Cowling 27.631.5041.35 Rocket Body 35.6600 4 Fins 74.016.981256.22 Total -20.481303.58 Xcp (Tail Reference in inches) 14.45

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Stability

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

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Fin Design

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Fin Specifications Dimensions based on flutter analysis, testing, and stability calculations: C r = 6” C t = 2.5” S = 4” t = 0.167” CtCt S CrCr

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Nose Cone Experiment

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Types of Nose Cones 1) Elliptical2) Conical http://myweb.cableone.net/cjcrowell/NCEQN2.DOC 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

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

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

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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.

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PROJECT COST MOTOR$1,227 PAYLOAD & RECOVERY$1345 NOSE & FINS$140 TOTAL COST$3,027 GIFTS IN KIND$555 TOTAL AMOUNT REQUIRED$2,477 CURRENT FUNDS$1,500 ADDITIONAL FUNDS REQURIED$977

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Launch Lug – helps to guide the rocket upward until it reaches enough velocity for the fins to engage. Parachute – assists in the safe recovery of the.

Launch Lug – helps to guide the rocket upward until it reaches enough velocity for the fins to engage. Parachute – assists in the safe recovery of the.

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