GOE 412 AIRFOIL (goe412-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 412 AIRFOIL (goe412-il) Reynolds number: 200,000 Max Cl/Cd: 82.91 at α=6.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe412-il-200000.txt Download as CSV file: xf-goe412-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 412 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.2137 0.09917 0.09566 -0.0686 0.9783 0.0707
-9.000 -0.2220 0.09283 0.08934 -0.0879 0.9683 0.0743
-8.750 -0.2046 0.08759 0.08410 -0.0877 0.9651 0.0752
-8.500 -0.1753 0.08496 0.08144 -0.0869 0.9637 0.0767
-8.250 -0.1515 0.08181 0.07826 -0.0898 0.9603 0.0788
-8.000 -0.1317 0.07770 0.07414 -0.0956 0.9566 0.0816
-7.750 -0.1596 0.06794 0.06413 -0.1185 0.9312 0.0867
-7.500 -0.1373 0.06449 0.06080 -0.1168 0.9289 0.0877
-7.250 -0.1217 0.06247 0.05878 -0.1154 0.9224 0.0886
-7.000 -0.1058 0.06018 0.05648 -0.1153 0.9157 0.0899
-6.750 -0.0873 0.05714 0.05337 -0.1175 0.9111 0.0922
-6.500 -0.0894 0.04957 0.04524 -0.1259 0.8984 0.1008
-6.250 -0.1033 0.03255 0.02695 -0.1268 0.8880 0.0626
-6.000 -0.0820 0.02919 0.02322 -0.1269 0.8836 0.0613
-5.750 -0.0605 0.02642 0.02001 -0.1266 0.8787 0.0605
-5.500 -0.0391 0.02443 0.01765 -0.1257 0.8720 0.0603
-5.250 -0.0114 0.02283 0.01571 -0.1257 0.8681 0.0612
-5.000 0.0155 0.02167 0.01425 -0.1255 0.8634 0.0624
-4.750 0.0403 0.02072 0.01307 -0.1249 0.8570 0.0631
-4.500 0.0693 0.01971 0.01185 -0.1249 0.8528 0.0639
-4.250 0.0974 0.01854 0.01063 -0.1249 0.8487 0.0655
-4.000 0.1218 0.01802 0.01011 -0.1243 0.8419 0.0674
-3.750 0.1505 0.01742 0.00945 -0.1242 0.8374 0.0700
-3.500 0.1801 0.01697 0.00886 -0.1243 0.8334 0.0739
-3.250 0.2035 0.01648 0.00846 -0.1236 0.8264 0.0790
-3.000 0.2319 0.01606 0.00797 -0.1235 0.8216 0.0867
-2.750 0.2605 0.01545 0.00747 -0.1235 0.8177 0.1068
-2.500 0.2846 0.01535 0.00745 -0.1229 0.8105 0.1333
-2.250 0.3127 0.01516 0.00725 -0.1229 0.8056 0.1530
-2.000 0.3425 0.01490 0.00698 -0.1231 0.8018 0.1726
-1.750 0.3658 0.01483 0.00701 -0.1225 0.7945 0.1928
-1.500 0.3932 0.01458 0.00689 -0.1224 0.7894 0.2222
-1.250 0.4225 0.01430 0.00677 -0.1226 0.7855 0.2803
-1.000 0.4448 0.01412 0.00691 -0.1219 0.7784 0.3594
-0.750 0.4678 0.01349 0.00692 -0.1210 0.7731 0.5309
-0.500 0.5356 0.01240 0.00685 -0.1280 0.7701 0.9935
-0.250 0.5659 0.01255 0.00690 -0.1287 0.7635 1.0000
0.000 0.5911 0.01263 0.00684 -0.1281 0.7573 1.0000
0.250 0.6193 0.01263 0.00668 -0.1279 0.7520 1.0000
0.500 0.6411 0.01274 0.00673 -0.1267 0.7431 1.0000
0.750 0.6695 0.01269 0.00653 -0.1265 0.7371 1.0000
1.000 0.6920 0.01282 0.00662 -0.1254 0.7290 1.0000
1.250 0.7191 0.01281 0.00650 -0.1250 0.7226 1.0000
1.500 0.7440 0.01290 0.00654 -0.1244 0.7156 1.0000
1.750 0.7693 0.01296 0.00654 -0.1238 0.7085 1.0000
2.000 0.7981 0.01299 0.00645 -0.1238 0.7033 1.0000
2.250 0.8204 0.01315 0.00663 -0.1228 0.6953 1.0000
2.500 0.8485 0.01316 0.00656 -0.1226 0.6893 1.0000
2.750 0.8724 0.01330 0.00669 -0.1219 0.6817 1.0000
3.000 0.8991 0.01332 0.00666 -0.1215 0.6745 1.0000
3.250 0.9240 0.01341 0.00672 -0.1209 0.6666 1.0000
3.500 0.9501 0.01344 0.00672 -0.1204 0.6586 1.0000
3.750 0.9751 0.01355 0.00682 -0.1198 0.6508 1.0000
4.000 1.0008 0.01363 0.00687 -0.1193 0.6430 1.0000
4.250 1.0258 0.01374 0.00698 -0.1187 0.6349 1.0000
4.500 1.0512 0.01379 0.00701 -0.1182 0.6259 1.0000
4.750 1.0746 0.01391 0.00715 -0.1173 0.6161 1.0000
5.000 1.1016 0.01397 0.00714 -0.1170 0.6076 1.0000
5.250 1.1230 0.01411 0.00734 -0.1158 0.5962 1.0000
5.500 1.1468 0.01425 0.00748 -0.1150 0.5861 1.0000
5.750 1.1716 0.01437 0.00754 -0.1144 0.5756 1.0000
6.000 1.1926 0.01453 0.00774 -0.1131 0.5625 1.0000
6.250 1.2137 0.01470 0.00792 -0.1118 0.5489 1.0000
6.500 1.2346 0.01491 0.00812 -0.1105 0.5350 1.0000
6.750 1.2552 0.01514 0.00836 -0.1091 0.5210 1.0000
7.000 1.2751 0.01541 0.00861 -0.1077 0.5066 1.0000
7.250 1.2941 0.01571 0.00890 -0.1061 0.4919 1.0000
7.500 1.3120 0.01605 0.00923 -0.1044 0.4765 1.0000
7.750 1.3287 0.01643 0.00959 -0.1025 0.4605 1.0000
8.000 1.3439 0.01686 0.01000 -0.1003 0.4439 1.0000
8.250 1.3570 0.01733 0.01044 -0.0978 0.4268 1.0000
8.500 1.3673 0.01786 0.01093 -0.0948 0.4095 1.0000
8.750 1.3765 0.01850 0.01152 -0.0918 0.3915 1.0000
9.000 1.3840 0.01925 0.01219 -0.0886 0.3725 1.0000
9.250 1.3908 0.02010 0.01297 -0.0855 0.3534 1.0000
9.500 1.3976 0.02101 0.01383 -0.0825 0.3344 1.0000
9.750 1.4041 0.02200 0.01478 -0.0797 0.3162 1.0000
10.000 1.4108 0.02306 0.01580 -0.0771 0.2996 1.0000
10.250 1.4181 0.02415 0.01687 -0.0747 0.2847 1.0000
10.500 1.4257 0.02528 0.01798 -0.0724 0.2715 1.0000
10.750 1.4330 0.02649 0.01916 -0.0702 0.2596 1.0000
11.000 1.4413 0.02767 0.02036 -0.0683 0.2482 1.0000
11.250 1.4502 0.02886 0.02160 -0.0666 0.2379 1.0000
11.500 1.4571 0.03023 0.02296 -0.0648 0.2281 1.0000
11.750 1.4633 0.03167 0.02441 -0.0630 0.2178 1.0000
12.000 1.4702 0.03313 0.02591 -0.0614 0.2073 1.0000
12.250 1.4750 0.03477 0.02757 -0.0598 0.1976 1.0000
12.500 1.4783 0.03658 0.02938 -0.0581 0.1881 1.0000
12.750 1.4844 0.03822 0.03110 -0.0568 0.1787 1.0000
13.000 1.4868 0.04022 0.03312 -0.0553 0.1695 1.0000
13.250 1.4883 0.04238 0.03531 -0.0539 0.1587 1.0000
13.500 1.4905 0.04458 0.03757 -0.0527 0.1447 1.0000
13.750 1.4916 0.04701 0.04002 -0.0518 0.1163 1.0000
14.000 1.4757 0.05118 0.04389 -0.0503 0.0870 1.0000
14.250 1.4599 0.05548 0.04811 -0.0489 0.0735 1.0000
14.500 1.4467 0.05970 0.05235 -0.0478 0.0613 1.0000
14.750 1.4376 0.06366 0.05631 -0.0470 0.0516 1.0000
15.000 1.4314 0.06738 0.06007 -0.0466 0.0457 1.0000
15.250 1.4233 0.07145 0.06420 -0.0463 0.0425 1.0000
15.500 1.4173 0.07536 0.06820 -0.0463 0.0399 1.0000
15.750 1.4087 0.07970 0.07260 -0.0464 0.0381 1.0000
16.000 1.3981 0.08436 0.07733 -0.0467 0.0367 1.0000
16.250 1.3930 0.08836 0.08146 -0.0471 0.0354 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 412 AIRFOIL (goe412-il)