GOE 10K AIRFOIL (goe10k-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 10K AIRFOIL (goe10k-il) Reynolds number: 100,000 Max Cl/Cd: 27.78 at α=3.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe10k-il-100000.txt Download as CSV file: xf-goe10k-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 10K AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.5449 0.10512 0.10016 -0.0120 1.0000 0.0551
-8.250 -0.5503 0.10288 0.09801 -0.0147 1.0000 0.0554
-8.000 -0.5529 0.10024 0.09544 -0.0184 1.0000 0.0555
-7.750 -0.5365 0.09376 0.08893 -0.0118 1.0000 0.0590
-7.500 -0.5370 0.09079 0.08601 -0.0124 1.0000 0.0606
-7.250 -0.5358 0.08743 0.08271 -0.0152 1.0000 0.0630
-7.000 -0.5321 0.08417 0.07945 -0.0215 1.0000 0.0645
-6.750 -0.5266 0.07986 0.07516 -0.0241 1.0000 0.0656
-6.500 -0.5201 0.07613 0.07145 -0.0213 1.0000 0.0688
-6.250 -0.5090 0.07248 0.06772 -0.0263 1.0000 0.0731
-6.000 -0.5006 0.06801 0.06323 -0.0274 1.0000 0.0756
-5.750 -0.4899 0.06455 0.05974 -0.0273 1.0000 0.0804
-5.500 -0.4754 0.06043 0.05545 -0.0305 1.0000 0.0856
-5.250 -0.4550 0.05910 0.05362 -0.0335 1.0000 0.0963
-5.000 -0.4497 0.05317 0.04807 -0.0304 1.0000 0.1009
-4.750 -0.4328 0.04965 0.04434 -0.0316 1.0000 0.1129
-4.500 -0.4165 0.04639 0.04092 -0.0316 1.0000 0.1259
-4.250 -0.4001 0.04316 0.03755 -0.0312 1.0000 0.1401
-4.000 -0.3835 0.04010 0.03441 -0.0304 1.0000 0.1568
-3.750 -0.3666 0.03727 0.03145 -0.0296 1.0000 0.1845
-3.500 -0.3510 0.03461 0.02877 -0.0281 1.0000 0.2175
-3.250 -0.2949 0.02814 0.02046 -0.0293 1.0000 0.0756
-3.000 -0.2659 0.02493 0.01639 -0.0273 1.0000 0.0551
-2.750 -0.2410 0.02232 0.01341 -0.0261 1.0000 0.0542
-2.500 -0.2148 0.01999 0.01069 -0.0249 1.0000 0.0528
-2.250 -0.1885 0.01819 0.00852 -0.0236 1.0000 0.0565
-2.000 -0.1633 0.01650 0.00680 -0.0225 1.0000 0.0665
-1.750 -0.1401 0.01505 0.00551 -0.0213 1.0000 0.0930
-1.500 -0.0810 0.01073 0.00387 -0.0268 1.0000 1.0000
-1.250 -0.0586 0.01076 0.00352 -0.0257 1.0000 1.0000
-1.000 -0.0362 0.01080 0.00327 -0.0246 1.0000 1.0000
-0.750 -0.0139 0.01086 0.00309 -0.0236 1.0000 1.0000
-0.500 0.0085 0.01093 0.00293 -0.0227 1.0000 1.0000
-0.250 0.0307 0.01101 0.00287 -0.0217 1.0000 1.0000
0.000 0.0529 0.01111 0.00286 -0.0208 1.0000 1.0000
0.250 0.0751 0.01122 0.00287 -0.0198 1.0000 1.0000
0.500 0.0973 0.01135 0.00294 -0.0190 1.0000 1.0000
0.750 0.1194 0.01149 0.00305 -0.0181 1.0000 1.0000
1.000 0.1415 0.01164 0.00321 -0.0172 1.0000 1.0000
1.250 0.1634 0.01181 0.00341 -0.0164 1.0000 1.0000
1.500 0.1853 0.01200 0.00364 -0.0155 1.0000 1.0000
1.750 0.2071 0.01220 0.00393 -0.0147 1.0000 1.0000
2.000 0.2287 0.01243 0.00429 -0.0138 1.0000 1.0000
2.250 0.2503 0.01267 0.00468 -0.0130 1.0000 1.0000
2.500 0.2716 0.01295 0.00512 -0.0122 1.0000 1.0000
2.750 0.2928 0.01324 0.00566 -0.0114 1.0000 1.0000
3.000 0.3138 0.01357 0.00624 -0.0106 1.0000 1.0000
3.250 0.5075 0.01827 0.00801 -0.0375 0.0560 1.0000
3.500 0.5335 0.02010 0.00995 -0.0362 0.0497 1.0000
3.750 0.5607 0.02260 0.01255 -0.0354 0.0449 1.0000
4.000 0.5887 0.02566 0.01590 -0.0344 0.0457 1.0000
4.250 0.6170 0.02795 0.01897 -0.0320 0.0520 1.0000
4.500 0.6418 0.03181 0.02300 -0.0308 0.0574 1.0000
5.250 0.7256 0.04518 0.03928 -0.0235 0.1758 1.0000
5.500 0.7394 0.04866 0.04276 -0.0217 0.1560 1.0000
6.250 0.7728 0.06009 0.05454 -0.0178 0.1134 1.0000
6.500 0.7772 0.06316 0.05802 -0.0173 0.1026 1.0000
6.750 0.7920 0.07126 0.06570 -0.0166 0.0982 1.0000
7.000 0.7878 0.07135 0.06651 -0.0170 0.0906 1.0000
7.250 0.8056 0.07934 0.07409 -0.0154 0.0854 1.0000
7.500 0.7929 0.08029 0.07558 -0.0172 0.0832 1.0000
7.750 0.7905 0.08422 0.07962 -0.0189 0.0796 1.0000
8.000 0.8049 0.08988 0.08514 -0.0160 0.0750 1.0000
8.250 0.7928 0.09333 0.08872 -0.0191 0.0745 1.0000
8.500 0.7809 0.09710 0.09250 -0.0218 0.0740 1.0000
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