GOE 391 AIRFOIL (goe391-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 391 AIRFOIL (goe391-il) Reynolds number: 100,000 Max Cl/Cd: 58.75 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe391-il-100000.txt Download as CSV file: xf-goe391-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 391 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
------ -------- --------- --------- -------- -------- --------
-8.000 -0.4325 0.09769 0.09268 -0.0217 1.0000 0.0483
-7.750 -0.4383 0.09603 0.09112 -0.0226 1.0000 0.0489
-7.500 -0.4395 0.09418 0.08934 -0.0259 1.0000 0.0494
-7.250 -0.3651 0.07705 0.07258 -0.0237 1.0000 0.0532
-7.000 -0.3691 0.07427 0.06986 -0.0226 1.0000 0.0542
-6.750 -0.3719 0.07134 0.06698 -0.0222 1.0000 0.0556
-6.500 -0.3743 0.06837 0.06406 -0.0222 1.0000 0.0570
-6.250 -0.3757 0.06537 0.06109 -0.0226 1.0000 0.0587
-6.000 -0.3754 0.06237 0.05809 -0.0238 1.0000 0.0606
-5.750 -0.3700 0.06007 0.05570 -0.0280 1.0000 0.0628
-5.500 -0.3591 0.05853 0.05388 -0.0321 1.0000 0.0637
-5.250 -0.3714 0.06635 0.06118 -0.0353 1.0000 0.0639
-5.000 -0.3733 0.06079 0.05591 -0.0314 1.0000 0.0663
-4.750 -0.3637 0.05795 0.05302 -0.0304 1.0000 0.0719
-4.500 -0.3398 0.05514 0.04987 -0.0338 1.0000 0.0786
-4.250 -0.3331 0.05162 0.04648 -0.0317 1.0000 0.0831
-4.000 -0.3085 0.04921 0.04367 -0.0337 1.0000 0.0923
-3.750 -0.2982 0.04579 0.04037 -0.0322 1.0000 0.0978
-3.500 -0.2785 0.04298 0.03735 -0.0326 1.0000 0.1085
-3.250 -0.2593 0.04043 0.03462 -0.0326 1.0000 0.1216
-3.000 -0.2414 0.03792 0.03201 -0.0321 1.0000 0.1371
-2.750 -0.2225 0.03568 0.02958 -0.0319 1.0000 0.1630
-2.500 -0.2074 0.03362 0.02751 -0.0309 1.0000 0.2045
-1.250 -0.0582 0.02343 0.01502 -0.0288 1.0000 0.1154
-1.000 -0.0306 0.02188 0.01304 -0.0277 1.0000 0.1002
-0.750 -0.0044 0.02088 0.01153 -0.0264 1.0000 0.0922
-0.500 0.0198 0.01979 0.01031 -0.0255 1.0000 0.0932
-0.250 0.0427 0.01903 0.00956 -0.0247 1.0000 0.0994
0.000 0.0658 0.01842 0.00888 -0.0238 1.0000 0.0996
0.250 0.0881 0.01795 0.00839 -0.0228 1.0000 0.1020
0.500 0.1105 0.01754 0.00801 -0.0219 1.0000 0.1062
0.750 0.1310 0.01735 0.00784 -0.0210 1.0000 0.1172
1.000 0.1535 0.01726 0.00775 -0.0204 1.0000 0.1303
1.250 0.1774 0.01715 0.00780 -0.0202 1.0000 0.1670
1.500 0.2342 0.01565 0.00799 -0.0263 0.9953 1.0000
1.750 0.2891 0.01632 0.00852 -0.0324 0.9830 1.0000
2.000 0.3409 0.01679 0.00893 -0.0377 0.9700 1.0000
2.250 0.3868 0.01711 0.00927 -0.0418 0.9573 1.0000
2.500 0.4350 0.01730 0.00955 -0.0460 0.9437 1.0000
2.750 0.4862 0.01730 0.00965 -0.0505 0.9293 1.0000
3.000 0.5329 0.01710 0.00960 -0.0538 0.9122 1.0000
3.250 0.6079 0.01564 0.00844 -0.0605 0.8866 1.0000
3.500 0.6539 0.01480 0.00782 -0.0623 0.8630 1.0000
3.750 0.6920 0.01416 0.00739 -0.0627 0.8366 1.0000
4.000 0.7266 0.01338 0.00674 -0.0617 0.7864 1.0000
4.250 0.7591 0.01292 0.00602 -0.0598 0.6633 1.0000
4.500 0.7784 0.01370 0.00606 -0.0567 0.5066 1.0000
4.750 0.7809 0.01591 0.00674 -0.0519 0.2401 1.0000
5.000 0.7906 0.01823 0.00807 -0.0487 0.1208 1.0000
5.250 0.8083 0.01950 0.00920 -0.0467 0.1044 1.0000
5.500 0.8271 0.02074 0.01037 -0.0450 0.0918 1.0000
5.750 0.8489 0.02209 0.01185 -0.0436 0.0850 1.0000
6.000 0.8727 0.02420 0.01382 -0.0428 0.0777 1.0000
6.250 0.8977 0.02572 0.01556 -0.0418 0.0704 1.0000
6.500 0.9254 0.02861 0.01847 -0.0415 0.0663 1.0000
6.750 0.9487 0.03097 0.02130 -0.0401 0.0622 1.0000
7.000 0.9704 0.03361 0.02441 -0.0384 0.0595 1.0000
7.250 0.9896 0.03721 0.02862 -0.0363 0.0601 1.0000
7.500 1.0049 0.04149 0.03343 -0.0340 0.0624 1.0000
7.750 1.0176 0.04634 0.03865 -0.0320 0.0643 1.0000
8.000 1.0299 0.05217 0.04473 -0.0305 0.0655 1.0000
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Polar data table (+)
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