MARSKE MONARCH AIRFOIL (NACA 43012A) (marske5-il)
MARSKE MONARCH AIRFOIL (NACA 43012A) - Marske Monarch airfoil (NACA 43012A)
Details | Dat file | Parser | |
(marske5-il) MARSKE MONARCH AIRFOIL (NACA 43012A) Marske Monarch airfoil (NACA 43012A) Max thickness 12.2% at 20% chord. Max camber 3.4% at 15% chord Source UIUC Airfoil Coordinates Database Source dat file The dat file is in Lednicer format |
MARSKE MONARCH AIRFOIL (NACA 43012A) 17. 17. 0.000000 0.000000 0.012500 0.038900 0.025000 0.051400 0.050000 0.069200 0.075000 0.080300 0.100000 0.087700 0.150000 0.093300 0.200000 0.092700 0.250000 0.089500 0.300000 0.085700 0.400000 0.077000 0.500000 0.065600 0.600000 0.052000 0.700000 0.037000 0.800000 0.024400 0.900000 0.012500 1.000000 0.000000 0.000000 0.000000 0.012500 -0.008300 0.025000 -0.011400 0.050000 -0.015400 0.075000 -0.018500 0.100000 -0.021100 0.150000 -0.025700 0.200000 -0.029500 0.250000 -0.032700 0.300000 -0.035000 0.400000 -0.039200 0.500000 -0.040300 0.600000 -0.039200 0.700000 -0.036000 0.800000 -0.027400 0.900000 -0.015000 1.000000 0.000000 |
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Polars for MARSKE MONARCH AIRFOIL (NACA 43012A) (marske5-il)
Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
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marske5-il | 50,000 | 9 | 15.7 at α=2° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
marske5-il | 50,000 | 5 | 20.8 at α=4.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
marske5-il | 100,000 | 9 | 23.3 at α=3.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
marske5-il | 100,000 | 5 | 31.8 at α=6° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
marske5-il | 200,000 | 9 | 36.1 at α=6.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
marske5-il | 200,000 | 5 | 47.7 at α=8° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
marske5-il | 500,000 | 9 | 65.1 at α=9° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
marske5-il | 500,000 | 5 | 73.6 at α=10.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
marske5-il | 1,000,000 | 9 | 92.9 at α=12.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
marske5-il | 1,000,000 | 5 | 96.5 at α=12° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
Reynolds number calculator |