GOE 654 AIRFOIL (goe654-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 654 AIRFOIL (goe654-il) Reynolds number: 50,000 Max Cl/Cd: 28.45 at α=4° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe654-il-50000-n5.txt Download as CSV file: xf-goe654-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 654 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
------ -------- --------- --------- -------- -------- --------
-9.500 -0.2283 0.11589 0.10958 -0.0374 1.0000 0.1432
-9.250 -0.2534 0.11588 0.10978 -0.0372 1.0000 0.1446
-9.000 -0.2779 0.11576 0.10986 -0.0354 1.0000 0.1449
-8.750 -0.2508 0.11086 0.10499 -0.0319 1.0000 0.1481
-8.500 -0.2255 0.10655 0.10066 -0.0360 0.9892 0.1522
-8.250 -0.2091 0.10267 0.09677 -0.0412 0.9757 0.1558
-7.750 -0.2178 0.08618 0.08013 -0.0643 0.9372 0.0839
-7.500 -0.1897 0.08227 0.07619 -0.0668 0.9288 0.0820
-7.250 -0.1731 0.07747 0.07135 -0.0723 0.9148 0.0807
-7.000 -0.1597 0.07243 0.06622 -0.0782 0.9005 0.0796
-6.750 -0.1480 0.06735 0.06102 -0.0837 0.8862 0.0787
-6.500 -0.1369 0.06231 0.05578 -0.0884 0.8725 0.0774
-6.250 -0.1253 0.05684 0.04996 -0.0929 0.8604 0.0754
-6.000 -0.1179 0.05158 0.04412 -0.0957 0.8464 0.0733
-5.750 -0.1052 0.04794 0.03984 -0.0965 0.8329 0.0722
-5.500 -0.0845 0.04522 0.03678 -0.0970 0.8219 0.0721
-5.250 -0.0612 0.04265 0.03381 -0.0976 0.8118 0.0720
-5.000 -0.0420 0.04063 0.03144 -0.0971 0.7999 0.0726
-4.750 -0.0158 0.03888 0.02954 -0.0975 0.7909 0.0742
-4.500 0.0064 0.03747 0.02790 -0.0971 0.7799 0.0757
-4.250 0.0305 0.03601 0.02613 -0.0967 0.7699 0.0769
-4.000 0.0577 0.03452 0.02428 -0.0966 0.7611 0.0776
-3.750 0.0814 0.03338 0.02285 -0.0960 0.7511 0.0785
-3.500 0.1108 0.03217 0.02134 -0.0959 0.7432 0.0797
-3.250 0.1343 0.03136 0.02030 -0.0950 0.7330 0.0810
-3.000 0.1649 0.03046 0.01909 -0.0949 0.7255 0.0836
-2.750 0.1870 0.02986 0.01849 -0.0938 0.7148 0.0867
-2.500 0.2183 0.02905 0.01752 -0.0938 0.7072 0.0912
-2.250 0.2403 0.02867 0.01701 -0.0926 0.6956 0.0949
-2.000 0.2682 0.02809 0.01626 -0.0922 0.6866 0.0993
-1.750 0.2930 0.02761 0.01571 -0.0914 0.6769 0.1060
-1.500 0.3173 0.02720 0.01526 -0.0906 0.6676 0.1178
-1.250 0.3438 0.02659 0.01476 -0.0902 0.6591 0.1467
-1.000 0.3653 0.02564 0.01455 -0.0895 0.6503 0.2816
-0.500 0.4584 0.02368 0.01436 -0.0941 0.6326 1.0000
-0.250 0.4834 0.02380 0.01413 -0.0933 0.6248 1.0000
0.000 0.5023 0.02418 0.01426 -0.0919 0.6155 1.0000
0.250 0.5283 0.02432 0.01409 -0.0913 0.6081 1.0000
0.500 0.5474 0.02474 0.01432 -0.0899 0.5988 1.0000
0.750 0.5739 0.02488 0.01419 -0.0894 0.5914 1.0000
1.000 0.5929 0.02534 0.01450 -0.0880 0.5820 1.0000
1.250 0.6197 0.02551 0.01443 -0.0876 0.5747 1.0000
1.500 0.6385 0.02602 0.01481 -0.0862 0.5654 1.0000
1.750 0.6652 0.02622 0.01480 -0.0857 0.5581 1.0000
2.000 0.6839 0.02676 0.01525 -0.0843 0.5489 1.0000
2.250 0.7099 0.02702 0.01533 -0.0838 0.5415 1.0000
2.500 0.7289 0.02759 0.01583 -0.0825 0.5327 1.0000
2.750 0.7533 0.02794 0.01603 -0.0818 0.5250 1.0000
3.000 0.7735 0.02849 0.01651 -0.0807 0.5169 1.0000
3.250 0.7951 0.02899 0.01692 -0.0797 0.5089 1.0000
3.500 0.8188 0.02942 0.01724 -0.0790 0.5020 1.0000
3.750 0.8347 0.03020 0.01802 -0.0775 0.4933 1.0000
4.000 0.8642 0.03038 0.01804 -0.0774 0.4877 1.0000
4.250 0.8729 0.03150 0.01924 -0.0751 0.4785 1.0000
4.500 0.8987 0.03185 0.01949 -0.0747 0.4725 1.0000
4.750 0.9111 0.03286 0.02053 -0.0729 0.4649 1.0000
5.000 0.9294 0.03358 0.02124 -0.0717 0.4583 1.0000
5.250 0.9583 0.03379 0.02134 -0.0716 0.4536 1.0000
5.500 0.9583 0.03543 0.02312 -0.0686 0.4452 1.0000
5.750 0.9828 0.03582 0.02346 -0.0681 0.4401 1.0000
6.000 0.9913 0.03709 0.02480 -0.0660 0.4337 1.0000
6.250 0.9996 0.03836 0.02612 -0.0640 0.4273 1.0000
6.500 1.0284 0.03853 0.02623 -0.0639 0.4232 1.0000
6.750 1.0151 0.04095 0.02878 -0.0599 0.4158 1.0000
7.000 1.0264 0.04200 0.02985 -0.0582 0.4104 1.0000
7.250 1.0616 0.04183 0.02964 -0.0587 0.4070 1.0000
7.500 1.0107 0.04674 0.03473 -0.0525 0.3980 1.0000
7.750 1.0326 0.04723 0.03523 -0.0518 0.3938 1.0000
8.000 1.0334 0.04931 0.03734 -0.0501 0.3884 1.0000
8.250 0.9963 0.05494 0.04309 -0.0480 0.3791 1.0000
8.500 1.0317 0.05414 0.04228 -0.0474 0.3766 1.0000
9.000 0.9579 0.06808 0.05639 -0.0468 0.3566 1.0000
9.500 0.9663 0.07340 0.06182 -0.0463 0.3451 1.0000
10.500 0.9281 0.09221 0.08085 -0.0484 0.3144 1.0000
10.750 0.9547 0.09185 0.08054 -0.0474 0.3122 1.0000
11.250 0.9365 0.10142 0.09023 -0.0490 0.2970 1.0000
11.500 0.9095 0.10900 0.09787 -0.0512 0.2874 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 654 AIRFOIL (goe654-il)