WORTMANN FX 082-512 AIRFOIL (fx082512-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: WORTMANN FX 082-512 AIRFOIL (fx082512-il) Reynolds number: 100,000 Max Cl/Cd: 46.6 at α=4° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-fx082512-il-100000.txt Download as CSV file: xf-fx082512-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: WORTMANN FX 082-512 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
------ -------- --------- --------- -------- -------- --------
-10.000 -0.6222 0.08357 0.07859 -0.0571 1.0000 0.0860
-9.750 -0.6260 0.07927 0.07429 -0.0562 1.0000 0.0841
-9.500 -0.6475 0.07432 0.06939 -0.0562 1.0000 0.0826
-9.250 -0.6808 0.06972 0.06483 -0.0560 1.0000 0.0811
-9.000 -0.7138 0.06225 0.05726 -0.0627 1.0000 0.0789
-8.750 -0.7274 0.05327 0.04781 -0.0748 1.0000 0.0772
-8.500 -0.7176 0.04752 0.04156 -0.0812 1.0000 0.0783
-8.250 -0.6969 0.04308 0.03655 -0.0865 1.0000 0.0817
-8.000 -0.6664 0.03980 0.03239 -0.0922 1.0000 0.0853
-7.750 -0.6429 0.03503 0.02752 -0.0945 1.0000 0.0901
-7.500 -0.6157 0.03281 0.02504 -0.0961 1.0000 0.0969
-7.250 -0.5871 0.03027 0.02219 -0.0978 1.0000 0.1029
-7.000 -0.5626 0.02851 0.02047 -0.0980 1.0000 0.1103
-6.750 -0.5358 0.02703 0.01882 -0.0986 1.0000 0.1171
-6.500 -0.5134 0.02588 0.01787 -0.0979 1.0000 0.1242
-6.250 -0.4873 0.02505 0.01698 -0.0980 1.0000 0.1309
-6.000 -0.4621 0.02432 0.01645 -0.0979 1.0000 0.1382
-5.750 -0.4321 0.02364 0.01576 -0.0992 1.0000 0.1463
-5.500 -0.3975 0.02276 0.01487 -0.1019 1.0000 0.1565
-5.250 -0.3548 0.02139 0.01384 -0.1071 1.0000 0.1901
-5.000 -0.3340 0.02228 0.01522 -0.1054 1.0000 0.2711
-4.750 -0.3131 0.02330 0.01605 -0.1037 1.0000 0.3038
-4.500 -0.2969 0.02471 0.01742 -0.1007 1.0000 0.3223
-4.250 -0.2825 0.02612 0.01885 -0.0973 1.0000 0.3360
-4.000 -0.2678 0.02741 0.02014 -0.0941 1.0000 0.3486
-3.750 -0.2478 0.02823 0.02083 -0.0926 1.0000 0.3610
-3.500 -0.2393 0.02945 0.02222 -0.0877 1.0000 0.3669
-3.250 -0.2157 0.02975 0.02237 -0.0876 1.0000 0.3772
-3.000 -0.2058 0.03072 0.02347 -0.0834 1.0000 0.3836
-2.750 -0.1847 0.03106 0.02372 -0.0827 1.0000 0.3938
-2.500 -0.1716 0.03173 0.02445 -0.0797 1.0000 0.4010
-2.250 -0.1473 0.03174 0.02437 -0.0801 1.0000 0.4095
-2.000 -0.1335 0.03226 0.02495 -0.0774 1.0000 0.4163
-1.750 -0.1025 0.03195 0.02446 -0.0800 1.0000 0.4238
-1.500 -0.0867 0.03217 0.02475 -0.0781 1.0000 0.4275
-1.250 -0.0635 0.03222 0.02476 -0.0784 1.0000 0.4333
-1.000 -0.0404 0.03231 0.02481 -0.0787 1.0000 0.4411
-0.750 -0.0112 0.03287 0.02539 -0.0796 0.9969 0.4501
-0.500 0.0294 0.03336 0.02589 -0.0826 0.9899 0.4586
-0.250 0.0723 0.03364 0.02612 -0.0865 0.9833 0.4661
0.000 0.1200 0.03360 0.02596 -0.0920 0.9777 0.4706
0.250 0.1567 0.03353 0.02590 -0.0947 0.9702 0.4726
0.500 0.1972 0.03354 0.02590 -0.0982 0.9637 0.4745
0.750 0.2392 0.03347 0.02581 -0.1021 0.9570 0.4758
1.000 0.2778 0.03342 0.02574 -0.1053 0.9492 0.4773
1.250 0.3310 0.03333 0.02563 -0.1108 0.9414 0.4797
1.500 0.3937 0.03274 0.02501 -0.1175 0.9277 0.4837
1.750 0.4547 0.03181 0.02413 -0.1231 0.9131 0.4868
2.000 0.4991 0.03112 0.02355 -0.1259 0.8996 0.4892
2.250 0.5660 0.02941 0.02197 -0.1314 0.8836 0.4928
2.500 0.6196 0.02805 0.02072 -0.1350 0.8693 0.4972
2.750 0.6672 0.02687 0.01964 -0.1376 0.8558 0.5016
3.000 0.7159 0.02524 0.01824 -0.1396 0.8411 0.5055
3.250 0.7752 0.02272 0.01595 -0.1425 0.8230 0.5102
3.500 0.8145 0.02127 0.01463 -0.1426 0.7959 0.5144
3.750 0.8644 0.01949 0.01295 -0.1440 0.7440 0.5192
4.000 0.9209 0.01976 0.01148 -0.1463 0.4171 0.5257
4.250 0.9243 0.02211 0.01245 -0.1423 0.2053 0.5293
4.500 0.9421 0.02330 0.01323 -0.1407 0.1661 0.5334
4.750 0.9594 0.02421 0.01409 -0.1387 0.1520 0.5373
5.000 0.9784 0.02507 0.01499 -0.1370 0.1424 0.5419
5.250 0.9982 0.02624 0.01609 -0.1357 0.1333 0.5471
5.750 1.0572 0.02926 0.01902 -0.1364 0.1134 0.5602
6.000 1.1002 0.03156 0.02129 -0.1390 0.1028 0.5691
6.250 1.1473 0.03483 0.02444 -0.1427 0.0921 0.5802
6.500 1.1748 0.03630 0.02634 -0.1418 0.0868 0.5891
6.750 1.2071 0.03893 0.02900 -0.1428 0.0801 0.6003
7.000 1.2349 0.04234 0.03281 -0.1424 0.0772 0.6128
7.250 1.2548 0.04424 0.03526 -0.1401 0.0749 0.6269
7.500 1.2729 0.04643 0.03797 -0.1379 0.0718 0.6440
7.750 1.2901 0.04935 0.04141 -0.1356 0.0707 0.6657
8.000 1.3038 0.05302 0.04572 -0.1327 0.0720 0.6962
8.250 1.3124 0.05688 0.05025 -0.1291 0.0741 0.7561
8.500 1.3121 0.05995 0.05382 -0.1242 0.0754 1.0000
10.750 0.8702 0.13810 0.13498 -0.1159 0.1780 0.6346
11.000 0.8942 0.13998 0.13703 -0.1130 0.1658 0.6621
|
Polar data table (+)
Polar graphs
<< Back to WORTMANN FX 082-512 AIRFOIL (fx082512-il)