Presentation on theme: "PANTHR Hybrid Rocket Final Design Review December 6 th 2006."— Presentation transcript:
PANTHR Hybrid Rocket Final Design Review December 6 th 2006
PANTHR Team Members Glen Guzik Niroshen Divitotawela Michael Harris Bruce Helming David Moschetti Danielle Pepe Jacob Teufert
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
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
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
OXIDIZER TANK FUEL GRAIN ABLATIVE LINER COMBUSTION CHAMBER NOZZLE INJECTOR MOTOR EXPLODED VIEW
Oxidizer Fill and Ignition System Fill internal oxidizer tank via external, commercial nitrous- oxide tank. Light solid propellant ignition charge via electric match.
Trajectory Analysis 1 Degree of Freedom Explicit First-Order Finite Difference Method Thrust and Mass=f(t) Drag=f(v) Density=f(h)
Regression Rate Use regression rate formula for hybrids a =.155, n=.5  Regression Rate = 1.98 mm/s  AA283 Aircraft and Rocket Propulsion – Hybrid Rockets. Stanford University Department of Aeronautics and Astronautics. 2004
Combustion Chamber Dimensioning From Trajectory Analysis: Average Mass Flow: 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
Combustion Chamber Thermodynamic Properties From Analysis Adiabatic Flame Temperature: 3800K From Literature Review Paraffin Flame Temperature: 1700K For Design Average Value: 2750K
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.
Specifications Conical Nozzle Ae/A* = 3.64 Divergence Angle of 8 o Length 3.87” Weight 1.13 lbs.
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 T Aluminum T Aluminum T Aluminum O Aluminum T 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.
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:
Max hoop stress (ANSYS) = 9640 psi Max hoop stress (Theory) = psi
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:
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:
Payload Data Collection Acceleration versus time in 3 dimensions Pressure versus time Flight video at 30 FPS 352 x 240
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
Stability Maintain the Static Margin Options: - Under-damped - Neutral - Over-damped Current Configuration: - over-damped _Stability.pdf
Stability Subsonic flight allows use of Barrowman Method Xcp (Tail reference) = inch Xcg (Tail reference) varies between – inches Center of Pressure X-bar (in)p(x)X*p(x) Nose Cone 326 Cowling Rocket Body Fins Total Xcp (Tail Reference in inches) 14.45
Fin Design 3 different fin designs based on initial rocket plans Flutter conditions accounted for Wind tunnel testing was performed
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
Nose Cone Experiment
Types of Nose Cones 1) Elliptical2) Conical They both have low drag characteristics in low- transonic Mach regions. Elliptical Shape Total Drag From Experiment = Small Length and Weight decrease Static Margin Conical Shape Total Drag From Experiment =0.041 Length and Weight increase Static Margin Final Choice: Elliptical
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
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
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.
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