GOE 565 AIRFOIL (goe565-il) Xfoil prediction polar at RE=100,000 Ncrit=9
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Airfoil: GOE 565 AIRFOIL (goe565-il) Reynolds number: 100,000 Max Cl/Cd: 56.93 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe565-il-100000.txt Download as CSV file: xf-goe565-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 565 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
------ -------- --------- --------- -------- -------- --------
-9.500 -0.4256 0.10644 0.10117 -0.0264 1.0000 0.0990
-9.250 -0.4430 0.10483 0.09969 -0.0309 1.0000 0.1013
-9.000 -0.4633 0.10309 0.09811 -0.0346 1.0000 0.1018
-8.750 -0.4245 0.09651 0.09140 -0.0285 1.0000 0.1059
-8.500 -0.3496 0.08417 0.07947 -0.0312 1.0000 0.1192
-8.250 -0.3421 0.08118 0.07650 -0.0298 1.0000 0.1250
-8.000 -0.3586 0.07854 0.07395 -0.0308 1.0000 0.1293
-7.750 -0.3868 0.07657 0.07214 -0.0315 1.0000 0.1304
-7.500 -0.3702 0.07177 0.06733 -0.0289 1.0000 0.1346
-7.250 -0.3705 0.06917 0.06476 -0.0268 1.0000 0.1390
-7.000 -0.3903 0.06703 0.06275 -0.0254 1.0000 0.1423
-6.750 -0.4310 0.06379 0.05955 -0.0322 1.0000 0.1456
-6.500 -0.4171 0.06024 0.05609 -0.0260 1.0000 0.1487
-6.250 -0.4608 0.06688 0.06234 -0.0307 1.0000 0.1498
-6.000 -0.4631 0.06297 0.05827 -0.0352 1.0000 0.1615
-5.750 -0.4579 0.06064 0.05574 -0.0366 1.0000 0.1751
-5.500 -0.4486 0.05778 0.05313 -0.0308 1.0000 0.1785
-5.250 -0.4412 0.05539 0.05072 -0.0297 1.0000 0.1934
-5.000 -0.4111 0.03967 0.03318 -0.0429 1.0000 0.1169
-4.750 -0.3893 0.03429 0.02706 -0.0427 1.0000 0.1014
-4.500 -0.3696 0.03117 0.02346 -0.0422 1.0000 0.1026
-4.250 -0.3477 0.02895 0.02079 -0.0414 1.0000 0.1023
-4.000 -0.3274 0.02673 0.01860 -0.0410 1.0000 0.1066
-3.750 -0.3046 0.02523 0.01679 -0.0402 1.0000 0.1094
-3.500 -0.2809 0.02382 0.01501 -0.0395 1.0000 0.1108
-3.250 -0.2579 0.02296 0.01380 -0.0386 1.0000 0.1150
-3.000 -0.2349 0.02179 0.01244 -0.0378 1.0000 0.1182
-2.750 -0.2124 0.02084 0.01144 -0.0370 1.0000 0.1213
-2.500 -0.1901 0.02016 0.01069 -0.0362 1.0000 0.1253
-2.250 -0.1681 0.01969 0.01007 -0.0353 1.0000 0.1317
-2.000 -0.1466 0.01900 0.00953 -0.0345 1.0000 0.1391
-1.750 -0.1251 0.01858 0.00909 -0.0335 1.0000 0.1465
-1.500 -0.1041 0.01811 0.00876 -0.0326 1.0000 0.1585
-1.250 -0.0829 0.01768 0.00853 -0.0319 1.0000 0.1832
-1.000 -0.0591 0.01633 0.00863 -0.0316 1.0000 0.4550
-0.750 -0.0119 0.01516 0.00882 -0.0347 0.9961 1.0000
-0.500 0.0328 0.01570 0.00906 -0.0386 0.9875 1.0000
-0.250 0.0769 0.01622 0.00934 -0.0423 0.9786 1.0000
0.000 0.1257 0.01675 0.00970 -0.0469 0.9699 1.0000
0.250 0.1695 0.01706 0.00986 -0.0503 0.9581 1.0000
0.500 0.2131 0.01732 0.01002 -0.0536 0.9459 1.0000
0.750 0.2562 0.01753 0.01016 -0.0567 0.9344 1.0000
1.000 0.3038 0.01770 0.01028 -0.0605 0.9248 1.0000
1.250 0.3459 0.01779 0.01035 -0.0632 0.9137 1.0000
1.500 0.3863 0.01778 0.01035 -0.0654 0.9010 1.0000
1.750 0.4314 0.01760 0.01019 -0.0682 0.8881 1.0000
2.000 0.4799 0.01725 0.00989 -0.0713 0.8756 1.0000
2.250 0.5262 0.01675 0.00945 -0.0737 0.8622 1.0000
2.500 0.5671 0.01626 0.00902 -0.0750 0.8475 1.0000
2.750 0.6043 0.01576 0.00860 -0.0754 0.8308 1.0000
3.000 0.6394 0.01531 0.00821 -0.0755 0.8134 1.0000
3.250 0.6732 0.01494 0.00790 -0.0754 0.7958 1.0000
3.500 0.7015 0.01472 0.00775 -0.0743 0.7731 1.0000
3.750 0.7324 0.01445 0.00750 -0.0736 0.7496 1.0000
4.000 0.7595 0.01437 0.00742 -0.0724 0.7214 1.0000
4.250 0.7849 0.01440 0.00746 -0.0710 0.6898 1.0000
4.500 0.8080 0.01451 0.00755 -0.0692 0.6518 1.0000
4.750 0.8300 0.01464 0.00756 -0.0672 0.6060 1.0000
5.000 0.8500 0.01493 0.00762 -0.0648 0.5518 1.0000
5.250 0.8683 0.01546 0.00788 -0.0625 0.4949 1.0000
5.500 0.8855 0.01610 0.00824 -0.0602 0.4382 1.0000
5.750 0.9033 0.01678 0.00871 -0.0581 0.3877 1.0000
6.000 0.9210 0.01752 0.00924 -0.0563 0.3428 1.0000
6.250 0.9389 0.01827 0.00986 -0.0545 0.3028 1.0000
6.500 0.9562 0.01901 0.01047 -0.0527 0.2611 1.0000
6.750 0.9705 0.02001 0.01122 -0.0506 0.2062 1.0000
7.000 0.9780 0.02208 0.01274 -0.0474 0.1360 1.0000
7.250 0.9931 0.02346 0.01389 -0.0453 0.1080 1.0000
7.500 1.0098 0.02488 0.01518 -0.0435 0.0954 1.0000
7.750 1.0287 0.02626 0.01660 -0.0419 0.0857 1.0000
8.000 1.0499 0.02811 0.01843 -0.0408 0.0792 1.0000
8.250 1.0727 0.02983 0.02022 -0.0399 0.0735 1.0000
8.500 1.0969 0.03225 0.02262 -0.0394 0.0683 1.0000
8.750 1.1200 0.03437 0.02509 -0.0384 0.0652 1.0000
9.000 1.1416 0.03676 0.02773 -0.0373 0.0625 1.0000
9.250 1.1612 0.03924 0.03027 -0.0365 0.0593 1.0000
9.500 1.1769 0.04308 0.03440 -0.0352 0.0574 1.0000
9.750 1.1864 0.04619 0.03801 -0.0328 0.0569 1.0000
10.000 1.1919 0.04979 0.04208 -0.0303 0.0567 1.0000
10.250 1.1937 0.05375 0.04646 -0.0277 0.0569 1.0000
10.500 1.1906 0.05770 0.05081 -0.0249 0.0569 1.0000
10.750 1.1833 0.06174 0.05520 -0.0220 0.0571 1.0000
11.000 1.1713 0.06568 0.05944 -0.0192 0.0571 1.0000
11.250 1.1549 0.06944 0.06343 -0.0161 0.0573 1.0000
11.500 1.1395 0.07383 0.06799 -0.0140 0.0576 1.0000
11.750 1.1248 0.07830 0.07265 -0.0127 0.0580 1.0000
12.000 1.0146 0.08907 0.08413 -0.0181 0.0644 1.0000
12.250 0.9768 0.09973 0.09488 -0.0253 0.0661 1.0000
12.500 0.9573 0.10897 0.10412 -0.0306 0.0680 1.0000
12.750 0.9168 0.12901 0.12409 -0.0421 0.0815 1.0000
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Polar data table (+)
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