GOE 7K AIRFOIL (goe07k-il)
GOE 7K AIRFOIL - Gottingen 7K airfoil
Details | Dat file | Parser | |
(goe07k-il) GOE 7K AIRFOIL Gottingen 7K airfoil Max thickness 11% at 50% chord. Max camber 4.5% at 50% chord Source UIUC Airfoil Coordinates Database Source dat file The dat file is in Lednicer format |
GOE 7K AIRFOIL 16. 16. 0.0000000 0.0000000 0.0250000 0.0231100 0.0500000 0.0344300 0.0750000 0.0433400 0.1000000 0.0513500 0.1500000 0.0641800 0.2000000 0.0741000 0.3000000 0.0891500 0.4000000 0.0975000 0.5000000 0.1004500 0.6000000 0.0980000 0.7000000 0.0871500 0.8000000 0.0686000 0.9000000 0.0409500 0.9500000 0.0237800 1.0000000 0.0000000 0.0000000 0.0000000 0.0250000 -.0102900 0.0500000 -.0116800 0.0750000 -.0124600 0.1000000 -.0129500 0.1500000 -.0136200 0.2000000 -.0133000 0.3000000 -.0120500 0.4000000 -.0108000 0.5000000 -.0095500 0.6000000 -.0083000 0.7000000 -.0070500 0.8000000 -.0058000 0.9000000 -.0045500 0.9500000 -.0039300 1.0000000 0.0000000 |
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Send to airfoil plotter Add to comparison Lednicer format dat file Selig format dat file |
Similar airfoils
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Polars for GOE 7K AIRFOIL (goe07k-il)
Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
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goe07k-il | 50,000 | 9 | 24.4 at α=9.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
goe07k-il | 50,000 | 5 | 25.2 at α=8.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
goe07k-il | 100,000 | 9 | 38.7 at α=7.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
goe07k-il | 100,000 | 5 | 40.7 at α=7.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
goe07k-il | 200,000 | 9 | 69.7 at α=6.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
goe07k-il | 200,000 | 5 | 70.2 at α=6° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
goe07k-il | 500,000 | 9 | 122.3 at α=4.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
goe07k-il | 500,000 | 5 | 116.5 at α=3.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
goe07k-il | 1,000,000 | 9 | 159.9 at α=2.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
goe07k-il | 1,000,000 | 5 | 142.8 at α=2.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
Reynolds number calculator |