GOE 9K AIRFOIL (goe09k-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 9K AIRFOIL (goe09k-il) Reynolds number: 50,000 Max Cl/Cd: 18.71 at α=2.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe09k-il-50000-n5.txt Download as CSV file: xf-goe09k-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 9K AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.6146 0.11736 0.11037 0.0112 1.0000 0.0735
-8.750 -0.6168 0.11512 0.10825 0.0082 1.0000 0.0755
-8.500 -0.6208 0.11318 0.10641 0.0046 1.0000 0.0761
-8.250 -0.6059 0.10692 0.10010 0.0071 1.0000 0.0799
-8.000 -0.6043 0.10389 0.09713 0.0055 1.0000 0.0832
-7.750 -0.6077 0.10189 0.09524 0.0002 1.0000 0.0857
-7.500 -0.6017 0.09726 0.09067 -0.0009 1.0000 0.0878
-7.250 -0.5928 0.09294 0.08635 0.0001 1.0000 0.0927
-7.000 -0.5879 0.09015 0.08358 -0.0065 1.0000 0.0983
-6.750 -0.5800 0.08549 0.07895 -0.0077 1.0000 0.1014
-6.500 -0.5704 0.08146 0.07494 -0.0076 1.0000 0.1081
-6.000 -0.5478 0.07404 0.06739 -0.0165 1.0000 0.1268
-5.750 -0.5353 0.07011 0.06341 -0.0181 1.0000 0.1402
-5.250 -0.5102 0.06189 0.05516 -0.0181 1.0000 0.1715
-4.750 -0.4379 0.05047 0.04276 -0.0270 1.0000 0.0561
-4.500 -0.4128 0.04623 0.03817 -0.0279 1.0000 0.0444
-4.250 -0.3868 0.04227 0.03375 -0.0284 1.0000 0.0380
-4.000 -0.3631 0.03861 0.02974 -0.0285 1.0000 0.0348
-3.750 -0.3349 0.03535 0.02577 -0.0282 1.0000 0.0314
-3.500 -0.3104 0.03220 0.02225 -0.0278 1.0000 0.0303
-3.250 -0.2843 0.02939 0.01891 -0.0271 1.0000 0.0298
-3.000 -0.2573 0.02688 0.01583 -0.0262 1.0000 0.0306
-2.750 -0.2301 0.02472 0.01313 -0.0251 1.0000 0.0317
-2.500 -0.2042 0.02250 0.01066 -0.0242 1.0000 0.0337
-2.250 -0.1774 0.02080 0.00859 -0.0229 1.0000 0.0384
-2.000 -0.1511 0.01935 0.00678 -0.0220 1.0000 0.0466
-1.750 -0.1252 0.01795 0.00523 -0.0212 1.0000 0.0684
-1.500 -0.0853 0.01345 0.00360 -0.0230 1.0000 1.0000
-1.250 -0.0611 0.01341 0.00292 -0.0221 1.0000 1.0000
-1.000 -0.0372 0.01338 0.00234 -0.0212 1.0000 1.0000
-0.750 -0.0133 0.01337 0.00199 -0.0204 1.0000 1.0000
-0.500 0.0104 0.01338 0.00174 -0.0197 1.0000 1.0000
-0.250 0.0341 0.01340 0.00158 -0.0189 1.0000 1.0000
0.000 0.0578 0.01343 0.00146 -0.0182 1.0000 1.0000
0.250 0.0814 0.01347 0.00144 -0.0175 1.0000 1.0000
0.500 0.1050 0.01352 0.00150 -0.0168 1.0000 1.0000
0.750 0.1284 0.01359 0.00165 -0.0160 1.0000 1.0000
1.000 0.1519 0.01367 0.00184 -0.0153 1.0000 1.0000
1.250 0.1752 0.01377 0.00211 -0.0146 1.0000 1.0000
1.500 0.1986 0.01388 0.00246 -0.0138 1.0000 1.0000
1.750 0.2219 0.01401 0.00301 -0.0131 1.0000 1.0000
2.000 0.2452 0.01417 0.00361 -0.0122 1.0000 1.0000
2.250 0.2685 0.01435 0.00433 -0.0114 1.0000 1.0000
2.750 0.3901 0.02152 0.00891 -0.0215 0.0365 1.0000
3.000 0.4187 0.02340 0.01104 -0.0208 0.0326 1.0000
3.250 0.4481 0.02555 0.01358 -0.0197 0.0299 1.0000
3.500 0.4765 0.02797 0.01650 -0.0186 0.0285 1.0000
3.750 0.5039 0.03066 0.01968 -0.0175 0.0289 1.0000
4.000 0.5298 0.03360 0.02311 -0.0163 0.0297 1.0000
4.250 0.5542 0.03675 0.02674 -0.0152 0.0305 1.0000
4.500 0.5769 0.04008 0.03052 -0.0142 0.0312 1.0000
4.750 0.5977 0.04377 0.03455 -0.0134 0.0320 1.0000
5.000 0.6216 0.04728 0.03880 -0.0124 0.0357 1.0000
5.250 0.6403 0.05139 0.04327 -0.0120 0.0382 1.0000
5.500 0.6603 0.05578 0.04808 -0.0118 0.0453 1.0000
6.000 0.7241 0.06869 0.06222 -0.0191 0.1785 1.0000
6.500 0.7408 0.07675 0.07038 -0.0198 0.1445 1.0000
6.750 0.7465 0.08101 0.07470 -0.0214 0.1326 1.0000
7.000 0.7599 0.08660 0.08021 -0.0189 0.1271 1.0000
7.250 0.7569 0.08958 0.08331 -0.0245 0.1177 1.0000
7.750 0.7594 0.09798 0.09168 -0.0298 0.1074 1.0000
8.000 0.7635 0.10225 0.09593 -0.0305 0.1031 1.0000
8.250 0.7748 0.10832 0.10197 -0.0286 0.0997 1.0000
8.500 0.7619 0.11070 0.10424 -0.0348 0.0980 1.0000
8.750 0.7577 0.11396 0.10744 -0.0382 0.0941 1.0000
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Polar data table (+)
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