GOE 602 MOD. AIRFOIL (goe602m-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 602 MOD. AIRFOIL (goe602m-il) Reynolds number: 100,000 Max Cl/Cd: 59.48 at α=6.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe602m-il-100000.txt Download as CSV file: xf-goe602m-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 602 MOD. 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 -0.3671 0.09437 0.08974 -0.0321 1.0000 0.1046
-7.750 -0.3836 0.09293 0.08841 -0.0312 1.0000 0.1071
-7.500 -0.4096 0.09214 0.08778 -0.0298 1.0000 0.1083
-7.250 -0.4373 0.09021 0.08598 -0.0358 1.0000 0.1093
-7.000 -0.4245 0.08609 0.08188 -0.0281 1.0000 0.1117
-6.750 -0.4185 0.08403 0.07984 -0.0243 1.0000 0.1165
-6.500 -0.4282 0.08146 0.07733 -0.0257 1.0000 0.1212
-6.250 -0.4415 0.07654 0.07239 -0.0361 1.0000 0.1248
-6.000 -0.4332 0.07540 0.07136 -0.0257 1.0000 0.1293
-5.500 -0.4300 0.06837 0.06433 -0.0298 1.0000 0.1422
-5.000 -0.4101 0.06138 0.05705 -0.0378 1.0000 0.1673
-4.750 -0.4038 0.05831 0.05410 -0.0345 1.0000 0.1704
-4.500 -0.3910 0.05537 0.05108 -0.0356 1.0000 0.1843
-4.250 -0.3286 0.03760 0.03158 -0.0527 1.0000 0.0982
-4.000 -0.3011 0.03343 0.02680 -0.0537 1.0000 0.0888
-3.750 -0.2710 0.02983 0.02206 -0.0542 1.0000 0.0820
-3.500 -0.2492 0.02788 0.01994 -0.0539 1.0000 0.0845
-3.250 -0.2205 0.02660 0.01841 -0.0547 0.9981 0.0887
-3.000 -0.1768 0.02508 0.01642 -0.0575 0.9921 0.0909
-2.750 -0.1339 0.02402 0.01504 -0.0604 0.9853 0.0970
-2.500 -0.0937 0.02311 0.01407 -0.0629 0.9778 0.1037
-2.250 -0.0502 0.02234 0.01323 -0.0659 0.9713 0.1118
-2.000 -0.0138 0.02161 0.01263 -0.0678 0.9622 0.1298
-1.750 0.0300 0.02013 0.01200 -0.0711 0.9569 0.2599
-1.500 0.0609 0.01953 0.01214 -0.0718 0.9467 0.4743
-1.250 0.0969 0.01941 0.01228 -0.0733 0.9374 0.5866
-1.000 0.1343 0.01872 0.01228 -0.0742 0.9304 0.7562
-0.750 0.1893 0.01835 0.01195 -0.0796 0.9215 1.0000
-0.500 0.2397 0.01845 0.01175 -0.0840 0.9148 1.0000
-0.250 0.2746 0.01853 0.01162 -0.0855 0.9031 1.0000
0.000 0.3116 0.01859 0.01152 -0.0873 0.8923 1.0000
0.250 0.3628 0.01843 0.01121 -0.0915 0.8864 1.0000
0.500 0.3971 0.01841 0.01108 -0.0925 0.8745 1.0000
0.750 0.4345 0.01831 0.01089 -0.0941 0.8638 1.0000
1.000 0.4858 0.01791 0.01043 -0.0979 0.8580 1.0000
1.250 0.5192 0.01778 0.01025 -0.0986 0.8458 1.0000
1.500 0.5538 0.01759 0.01003 -0.0993 0.8341 1.0000
1.750 0.5905 0.01732 0.00973 -0.1003 0.8230 1.0000
2.000 0.6341 0.01682 0.00923 -0.1023 0.8146 1.0000
2.250 0.6663 0.01660 0.00899 -0.1024 0.8010 1.0000
2.500 0.6986 0.01638 0.00876 -0.1025 0.7870 1.0000
2.750 0.7307 0.01617 0.00856 -0.1026 0.7725 1.0000
3.000 0.7611 0.01604 0.00841 -0.1023 0.7568 1.0000
3.250 0.7883 0.01606 0.00842 -0.1016 0.7393 1.0000
3.500 0.8165 0.01609 0.00847 -0.1011 0.7221 1.0000
3.750 0.8449 0.01613 0.00849 -0.1006 0.7046 1.0000
4.000 0.8737 0.01618 0.00851 -0.1002 0.6869 1.0000
4.250 0.8985 0.01637 0.00870 -0.0991 0.6670 1.0000
4.500 0.9238 0.01655 0.00890 -0.0981 0.6472 1.0000
4.750 0.9507 0.01673 0.00902 -0.0974 0.6282 1.0000
5.000 0.9731 0.01702 0.00936 -0.0961 0.6074 1.0000
5.250 0.9975 0.01727 0.00960 -0.0950 0.5873 1.0000
5.500 1.0198 0.01753 0.00991 -0.0935 0.5650 1.0000
5.750 1.0400 0.01768 0.01001 -0.0915 0.5373 1.0000
6.000 1.0577 0.01784 0.01012 -0.0891 0.5056 1.0000
6.250 1.0748 0.01807 0.01030 -0.0868 0.4737 1.0000
6.500 1.0936 0.01843 0.01068 -0.0849 0.4474 1.0000
6.750 1.1091 0.01879 0.01101 -0.0824 0.4132 1.0000
7.000 1.1233 0.01923 0.01141 -0.0798 0.3750 1.0000
7.250 1.1371 0.01978 0.01189 -0.0772 0.3338 1.0000
7.500 1.1484 0.02055 0.01249 -0.0743 0.2826 1.0000
7.750 1.1521 0.02199 0.01339 -0.0706 0.1952 1.0000
8.000 1.1485 0.02442 0.01496 -0.0663 0.0975 1.0000
8.250 1.1520 0.02628 0.01664 -0.0627 0.0690 1.0000
8.500 1.1583 0.02768 0.01806 -0.0593 0.0589 1.0000
8.750 1.1627 0.02917 0.01956 -0.0559 0.0544 1.0000
9.000 1.1682 0.03059 0.02113 -0.0527 0.0515 1.0000
9.250 1.1730 0.03211 0.02274 -0.0497 0.0489 1.0000
9.500 1.1765 0.03381 0.02448 -0.0469 0.0466 1.0000
9.750 1.1805 0.03582 0.02650 -0.0442 0.0446 1.0000
10.000 1.1920 0.03758 0.02838 -0.0422 0.0429 1.0000
10.250 1.2107 0.03956 0.03045 -0.0409 0.0421 1.0000
10.500 1.2389 0.04194 0.03299 -0.0405 0.0413 1.0000
10.750 1.2728 0.04504 0.03634 -0.0409 0.0411 1.0000
11.000 1.2988 0.04871 0.04036 -0.0407 0.0414 1.0000
11.250 1.3086 0.05213 0.04416 -0.0388 0.0415 1.0000
11.500 1.3098 0.05537 0.04776 -0.0362 0.0415 1.0000
11.750 1.3046 0.05858 0.05130 -0.0334 0.0413 1.0000
12.000 1.2954 0.06198 0.05503 -0.0307 0.0413 1.0000
12.250 1.2830 0.06572 0.05908 -0.0284 0.0416 1.0000
12.500 1.2677 0.06974 0.06339 -0.0267 0.0419 1.0000
12.750 1.2501 0.07413 0.06805 -0.0256 0.0423 1.0000
13.000 1.2311 0.07893 0.07310 -0.0253 0.0427 1.0000
13.250 1.2102 0.08415 0.07855 -0.0260 0.0430 1.0000
13.500 1.1890 0.08990 0.08450 -0.0275 0.0435 1.0000
13.750 1.1668 0.09617 0.09096 -0.0299 0.0438 1.0000
14.000 1.1460 0.10297 0.09791 -0.0330 0.0443 1.0000
14.250 1.1266 0.11020 0.10526 -0.0366 0.0447 1.0000
14.500 1.1132 0.11715 0.11233 -0.0399 0.0454 1.0000
14.750 0.8234 0.14629 0.14221 -0.0614 0.0527 1.0000
15.000 0.8192 0.15154 0.14746 -0.0636 0.0543 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 602 MOD. AIRFOIL (goe602m-il)