GOE 460 AIRFOIL (goe460-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 460 AIRFOIL (goe460-il) Reynolds number: 50,000 Max Cl/Cd: 20.79 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe460-il-50000.txt Download as CSV file: xf-goe460-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 460 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.000 -0.2982 0.12267 0.11336 -0.0184 1.0000 0.4841
-10.750 -0.2755 0.11935 0.11001 -0.0187 1.0000 0.4919
-10.500 -0.2992 0.11916 0.10987 -0.0165 1.0000 0.5009
-10.250 -0.2706 0.11479 0.10547 -0.0174 1.0000 0.5071
-10.000 -0.3613 0.10474 0.09541 -0.0206 1.0000 0.4225
-9.750 -0.3604 0.10095 0.09162 -0.0205 1.0000 0.4195
-9.500 -0.3710 0.09700 0.08767 -0.0201 1.0000 0.4178
-9.250 -0.3894 0.09312 0.08381 -0.0191 1.0000 0.4199
-9.000 -0.4244 0.08835 0.07909 -0.0176 1.0000 0.4228
-8.750 -0.3907 0.08846 0.07919 -0.0161 1.0000 0.4334
-8.500 -0.4108 0.08510 0.07586 -0.0142 1.0000 0.4384
-8.250 -0.4622 0.07973 0.07057 -0.0112 1.0000 0.4422
-8.000 -0.4190 0.08093 0.07174 -0.0099 1.0000 0.4523
-7.750 -0.4505 0.07738 0.06825 -0.0067 1.0000 0.4584
-7.500 -0.5754 0.06724 0.05827 0.0000 1.0000 0.4625
-7.250 -0.7731 0.05462 0.04584 0.0172 1.0000 0.4646
-7.000 -0.5731 0.06548 0.05655 0.0063 1.0000 0.4804
-6.750 -0.8460 0.04951 0.04071 0.0320 1.0000 0.4797
-6.500 -0.6417 0.05914 0.05032 0.0167 1.0000 0.4973
-6.250 -0.8631 0.04711 0.03828 0.0414 1.0000 0.4992
-6.000 -0.7941 0.04994 0.04121 0.0381 1.0000 0.5117
-5.750 -0.8870 0.04431 0.03540 0.0517 1.0000 0.5207
-5.500 -0.8262 0.04659 0.03783 0.0487 1.0000 0.5321
-5.250 -0.8598 0.04389 0.03504 0.0560 1.0000 0.5434
-5.000 -0.8435 0.04377 0.03494 0.0577 1.0000 0.5545
-4.750 -0.8412 0.04291 0.03407 0.0610 1.0000 0.5666
-4.500 -0.8563 0.04125 0.03233 0.0663 1.0000 0.5813
-4.250 -0.8283 0.04191 0.03306 0.0667 1.0000 0.5945
-4.000 -0.8145 0.04184 0.03301 0.0687 1.0000 0.6085
-3.750 -0.8120 0.04119 0.03234 0.0721 1.0000 0.6241
-3.500 -0.8119 0.04049 0.03160 0.0757 1.0000 0.6407
-3.250 -0.7815 0.04153 0.03269 0.0757 1.0000 0.6546
-3.000 -0.7589 0.04220 0.03339 0.0765 1.0000 0.6693
-2.750 -0.7462 0.04237 0.03355 0.0786 1.0000 0.6850
-2.500 -0.6693 0.04489 0.03601 0.0700 0.9844 0.7032
-2.250 -0.6108 0.04652 0.03757 0.0642 0.9703 0.7205
-2.000 -0.5317 0.04901 0.03999 0.0558 0.9562 0.7349
-1.750 -0.4185 0.05228 0.04321 0.0423 0.9417 0.7459
-1.500 -0.3556 0.05352 0.04438 0.0358 0.9294 0.7614
-1.250 -0.3305 0.05371 0.04454 0.0350 0.9170 0.7781
-1.000 -0.2163 0.05628 0.04709 0.0215 0.9027 0.7902
-0.750 -0.1457 0.05733 0.04812 0.0139 0.8902 0.8071
-0.500 -0.0776 0.05787 0.04865 0.0065 0.8780 0.8245
-0.250 -0.0674 0.05787 0.04864 0.0077 0.8649 0.8393
0.000 0.0000 0.05811 0.04888 0.0000 0.8522 0.8523
0.250 0.0673 0.05787 0.04864 -0.0077 0.8393 0.8649
0.500 0.0772 0.05787 0.04864 -0.0065 0.8244 0.8779
0.750 0.1450 0.05734 0.04813 -0.0139 0.8073 0.8902
1.000 0.2162 0.05628 0.04709 -0.0215 0.7904 0.9027
1.250 0.3303 0.05370 0.04453 -0.0350 0.7781 0.9170
1.500 0.3558 0.05350 0.04437 -0.0358 0.7614 0.9295
1.750 0.4197 0.05222 0.04314 -0.0425 0.7458 0.9418
2.000 0.5307 0.04902 0.04000 -0.0557 0.7346 0.9563
2.250 0.6115 0.04648 0.03753 -0.0643 0.7205 0.9705
2.500 0.6683 0.04493 0.03605 -0.0699 0.7033 0.9844
2.750 0.7443 0.04244 0.03362 -0.0783 0.6852 0.9998
3.000 0.7587 0.04219 0.03339 -0.0765 0.6694 1.0000
3.250 0.7819 0.04150 0.03266 -0.0758 0.6549 1.0000
3.500 0.8116 0.04049 0.03159 -0.0757 0.6408 1.0000
3.750 0.8123 0.04115 0.03229 -0.0721 0.6241 1.0000
4.000 0.8145 0.04182 0.03299 -0.0687 0.6085 1.0000
4.250 0.8284 0.04188 0.03303 -0.0667 0.5945 1.0000
4.500 0.8566 0.04120 0.03227 -0.0664 0.5812 1.0000
4.750 0.8406 0.04293 0.03408 -0.0609 0.5667 1.0000
5.000 0.8437 0.04373 0.03490 -0.0577 0.5546 1.0000
5.250 0.8605 0.04382 0.03497 -0.0561 0.5433 1.0000
5.500 0.8262 0.04658 0.03782 -0.0487 0.5323 1.0000
5.750 0.8872 0.04430 0.03538 -0.0517 0.5208 1.0000
6.000 0.7944 0.04991 0.04118 -0.0382 0.5118 1.0000
6.250 0.8639 0.04706 0.03823 -0.0415 0.4993 1.0000
6.500 0.6487 0.05872 0.04990 -0.0172 0.4972 1.0000
6.750 0.8453 0.04950 0.04070 -0.0319 0.4796 1.0000
7.000 0.5770 0.06516 0.05623 -0.0065 0.4804 1.0000
7.250 0.4591 0.07422 0.06515 0.0011 0.4752 1.0000
7.500 0.5795 0.06689 0.05793 -0.0002 0.4626 1.0000
7.750 0.4513 0.07728 0.06814 0.0067 0.4584 1.0000
8.000 0.4184 0.08095 0.07175 0.0099 0.4527 1.0000
8.250 0.4636 0.07961 0.07044 0.0112 0.4423 1.0000
8.500 0.4116 0.08500 0.07576 0.0142 0.4385 1.0000
8.750 0.3923 0.08827 0.07899 0.0162 0.4329 1.0000
9.000 0.4250 0.08828 0.07901 0.0176 0.4228 1.0000
9.250 0.3891 0.09303 0.08372 0.0192 0.4195 1.0000
9.500 0.3713 0.09691 0.08758 0.0201 0.4175 1.0000
9.750 0.3605 0.10086 0.09152 0.0205 0.4192 1.0000
10.000 0.3600 0.10467 0.09534 0.0206 0.4222 1.0000
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Polar data table (+)
Polar graphs
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