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									                       PANTHR
                      Hybrid Rocket




Final Design Review
 December 6th 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       • Payload and Recovery

  - Niroshen Divitotawela     - Glen Guzik
                              - Bruce Helming
  - Michael Harris
                              - Danielle Pepe
  - Jacob Teufert
                              - Michael Harris

• Aerodynamics and Flight   • Structural Analysis
  Stability
                              - David Moschetti
  - Bruce Helming             - Niroshen Divitotawela
  - Danielle Pepe
                            • Safety and Logistics
                              -David Moschetti
                              -Glen Guzik
    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
          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
                 MOTOR
                   EXPLODED VIEW




                                        OXIDIZER TANK

         ABLATIVE LINER


                                   INJECTOR



                     FUEL GRAIN


NOZZLE
                COMBUSTION CHAMBER
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
                                                           r  aG
                                                                       n
                                                                        ox
     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
           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
    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”
       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                                                                1
                                                  1
                                                                      2
  - Decided to expand the         2  P
                            Me            0
                                                   
                                                  
                                                                
                                                              1
    flow to sea-level                1   Pe
                                         
                                                  
                                                               
    pressure.                    
                                                              
  - Use of isentropic
    relations                                                  1
                            Ae 1  2    1              2
                                                              2 1
  - Find the Area Ratio               1    M e 
  - From trajectory         A M e  1 
                             
                                             2  
    computation make use
    of estimate of mass                           1
    flow rate.
                               m RT0  2 
                                               2 1
                            A        
                                        1
                                 P0       
                                            
                       Non-Ideal Expansion
                                   Rocket Launch Ae/A*=3.64

          1.400
                                                   Very Small losses     0 ft
                                                   associated with
                      Designed                                           2000 ft
                                                   non-ideal expansion
                      Ae/A*=3.64                                         4000 ft
                                                   for design altitude
          1.200
                                                                         6000 ft
                                                                         8000 ft
                                                                         10000 ft
          1.000                                                          12000 ft
                                                                         Designed Ae/A*


          0.800
T/Tconv




          0.600




          0.400




          0.200




          0.000
              1.000                       10.000                                    100.000
                                           Ae/A*
              Specifications

•   Conical Nozzle
•   Ae/A* = 3.64
•   Divergence Angle of 8o
•   Length 3.87”
•   Weight 1.13 lbs.
                                      Trade Study
                         Strength/
                            Density                                 Corrosion
      Material             Ratio      Weldability   Machinability   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             Several Alloys were compared in the decision
   Below Average     1            process for the material of the tubing needed for the
   Average           2            tank.
   Above Average     3
                                   Al 6061-T6 was observed to be the best metal to use
   Far Above Average 4
                                  considering cost and strength. Ratings were acquired by
                                  the Hadco Aluminum website.
             Structural Analysis
• Most severely                      pr                pr
  stressed components       1               t
  are the Combustion                  t               1
  Chamber and
  Oxidizer Tank            750 psi *1.5in. * F .S .
                        t                           0.1125 in.
                                20000 psi
• Wall Thickness was
  calculated using hoop            With F.S. of 2:
  stress equation
• Max hoop stress (ANSYS) = 9640 psi
• Max hoop stress (Theory) = 10000 psi
            Structural Analysis
                        Total Force acting on Bulkheads:
• We are using 8
  bolts the attach     F  PA  750 psi * 5.94 in.2  4455 lb.
  each bulkhead
• Each bolt is          Shear on each Bolt:
  made of 1022
  Carbon Steel              F * F .S .     4455 lb. * 2
                                                       22 ,688 psi
• The Allowable
                                                     2
                             ABolt Bi    0.04909 in. * 8
  Shear for each
  bolt is 29,000 psi
            Structural Analysis
• The Bearing Stress was        Bearing Stress Yield:
  calculated for the           b , yield  1.5 * Tensileyield
  aluminum tube using the
  force of the load                     1.5 * 35000 psi
  distributed to each bolt     b , yield  52500 psi
• With that the calculation
  divides the load by the       Total Bearing Stress:
  thickness of the wall,
  diameter of each hole,           F * F .S .
                              b 
  and the number of bolts             td
• The allowable was found                   4455lb. * 2
  to be 1.5 times the            
                                   0.125in. * 0.25in. * (8 bolts)
  allowable Tensile
  strength                     b  35640 psi
Payload Layout
            Payload Data Collection




• Acceleration versus time in 3 dimensions
• Pressure versus time
• Flight video at 30 FPS 352 x 240
                                            Pressure Distribution Behind Various Rocket Bodies

                            1.5



                           1.45
Dynamic Pressure(in H2O)




                            1.4



                           1.35



                            1.3



                           1.25                                                              Cowling Configuration
                                                                                             Uniform Diameter
                                                                                             Smooth Transition
                            1.2                                                              Poly. (Cowling Configuration)
                                                                                             Poly. (Uniform Diameter)
                                                                                             Poly. (Smooth Transition)
                           1.15
                                  35   45                55               65            75                        85
                                                              Probe Position (mm)
    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
                        http://www.rockets4schools.org/education/Rocket
                        _Stability.pdf
                     Stability
• Subsonic flight allows                     cg 
                                                   x W
  use of Barrowman                                W
  Method
• Xcp (Tail reference)                       cp 
                                                   x  p( x)dx
      = 14.45 inch                                    p ( x)
• Xcg (Tail reference)                                      Center of Pressure

  varies between
                                                 X-bar (in)         p(x)       X*p(x)
  34.26 – 38.25 inches     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)             14.45
                   Stability
                    CP and CG location




              72




              60




              48

                                         Center of
Height (in)




                                         Pressure

              36
                                         Center of
                                         Gravity



              24




              12




              0
Stability
                Fin Design

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


       1.05




         1




       0.95
q/q∞




        0.9




                                                                                                   Tapered Sweep
       0.85
                                                                                                   Clipped Delta

                                                                                                   Trapezoidal

        0.8
          -1.000   -0.800   -0.600   -0.400     -0.200      0.000      0.200       0.400   0.600      0.800        1.000
                                                              y/c
           Fin Specifications
• Dimensions
  based on flutter
  analysis, testing,    Ct
  and stability
  calculations:
• Cr = 6”
• Ct = 2.5”                     S
• S = 4”
• t = 0.167”
                        Cr
                       Nose Cone Experiment
                                 Wake Profile Pressure Distribution Vs. Transversing Distance


                                                         1.2




                                                           1




                                                         0.8
Pressure q/q'




                                                         0.6




                                                         0.4




                                                         0.2


                                                                                     Ellipse         Cone

                                                           0
                -0.5   -0.4   -0.3     -0.2      -0.1          0       0.1     0.2             0.3      0.4   0.5
                                                 Transversing Distance y/d
                              Types of Nose Cones
    1) Elliptical                                      2) Conical




                      •They both have low drag characteristics in low-
                      transonic Mach regions.

   Elliptical Shape                                  Conical Shape
   •Total Drag From Experiment = 0.029               •Total Drag From Experiment =0.041
   •Small Length and Weight decrease                 •Length and Weight increase Static Margin
   Static Margin

                                       Final Choice: Elliptical


http://myweb.cableone.net/cjcrowell/NCEQN2.DOC
                                         Nosecone
                Recovery
• Barometric Altimeter
• Drogue Chute –
  Deploys at apogee                    Drogue Parachute
• Main Chute – Deploys
  when altimeter
  detects specified
  altitude (~1500ft)


                     Main Parachute


                                      Cut Away View
     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
                   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
Questions?

								
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