PANTHR Hybrid Rocket

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PANTHRHybrid Rocket
Final Design ReviewDecember 6th 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 (CnHm
  n25, m50)
• - 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

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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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MOTOR EXPLODED VIEW
OXIDIZER TANK
ABLATIVE LINER
INJECTOR
FUEL GRAIN
NOZZLE
COMBUSTION CHAMBER
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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 Massf(t)
• Dragf(v)
• Densityf(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 (drtb)
• 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
Combustion Chamber
Ablative Liner
4.2
Fuel Grain
1.14
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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 8o
• Length 3.87
• Weight 1.13 lbs.

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Trade Study
Material Strength/ Density Ratio Weldability Machinability Corrosion Resistance Availability Cost Score
Aluminum 7075 - T6 4 1 3 2 1 2 13
Aluminum 2024 - T3 3.5 2 3 1 3 1 13.5
Aluminum 6061 - T6 3 2 2 3 4 4 18
Aluminum 6061 - O 1 4 1 3 1 1 11
Aluminum 6061 - T4 2.5 3 2 3 1 1 12.5
Scale
Far Below Average 0
Below Average 1
Average 2
Above Average 3
Far Above Average 4

• 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
Total Force acting on Bulkheads

• 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
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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
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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_St
ability.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 Center of Pressure Center of Pressure
X-bar (in) p(x) Xp(x)
Nose Cone 3 2 6
Cowling 27.63 1.50 41.35
Rocket Body 35.66 0 0
4 Fins 74.0 16.98 1256.22
Total - 20.48 1303.58
Xcp (Tail Reference in inches) Xcp (Tail Reference in inches) 14.45
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Stability
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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
• Cr 6
• Ct 2.5
• S 4
• t 0.167

Ct
S
Cr
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Nose Cone Experiment
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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
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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
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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

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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 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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Questions?