EPPLER 417 AIRFOIL (e417-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 417 AIRFOIL (e417-il) Reynolds number: 50,000 Max Cl/Cd: 33.17 at α=8° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e417-il-50000.txt Download as CSV file: xf-e417-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 417 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.5345 0.10421 0.09775 -0.0345 1.0000 0.1732
-8.750 -0.5876 0.09419 0.08792 -0.0406 1.0000 0.1527
-8.500 -0.6432 0.08895 0.08281 -0.0415 1.0000 0.1513
-8.250 -0.6928 0.08427 0.07812 -0.0426 1.0000 0.1504
-8.000 -0.7069 0.07784 0.07156 -0.0433 1.0000 0.1372
-7.750 -0.7133 0.07266 0.06622 -0.0433 1.0000 0.1279
-7.500 -0.7333 0.06649 0.05942 -0.0450 1.0000 0.1186
-7.250 -0.7268 0.06170 0.05444 -0.0448 1.0000 0.1132
-7.000 -0.7216 0.05655 0.04827 -0.0463 1.0000 0.1051
-6.750 -0.7054 0.05249 0.04400 -0.0461 1.0000 0.1025
-6.500 -0.6869 0.04875 0.03983 -0.0462 1.0000 0.0998
-6.250 -0.6653 0.04541 0.03598 -0.0462 1.0000 0.0976
-6.000 -0.6423 0.04261 0.03274 -0.0458 1.0000 0.0974
-5.750 -0.6195 0.04043 0.03025 -0.0451 1.0000 0.0995
-5.500 -0.5965 0.03858 0.02806 -0.0441 1.0000 0.1027
-5.250 -0.5734 0.03707 0.02620 -0.0425 1.0000 0.1058
-5.000 -0.5554 0.03563 0.02494 -0.0393 1.0000 0.1097
-4.750 -0.5376 0.03472 0.02405 -0.0359 1.0000 0.1159
-4.500 -0.5211 0.03376 0.02317 -0.0328 1.0000 0.1254
-4.250 -0.5031 0.03246 0.02201 -0.0311 1.0000 0.1460
-4.000 -0.4840 0.02802 0.02059 -0.0320 1.0000 0.4067
-3.750 -0.5010 0.03407 0.02685 -0.0129 1.0000 0.6319
-3.500 -0.5124 0.03730 0.03006 0.0016 1.0000 0.6788
-3.250 -0.5247 0.04006 0.03283 0.0172 1.0000 0.7302
-3.000 -0.5221 0.04221 0.03486 0.0311 1.0000 0.7881
-2.750 -0.4665 0.04334 0.03544 0.0330 1.0000 0.8472
-2.500 -0.4116 0.04289 0.03453 0.0285 1.0000 0.8743
-2.250 -0.3831 0.04210 0.03347 0.0271 1.0000 0.8891
-2.000 -0.3600 0.04130 0.03244 0.0263 1.0000 0.9004
-1.750 -0.3620 0.04042 0.03146 0.0297 1.0000 0.9088
-1.500 -0.3304 0.03978 0.03060 0.0272 1.0000 0.9172
-1.250 -0.3238 0.03903 0.02974 0.0290 1.0000 0.9240
-1.000 -0.3018 0.03846 0.02900 0.0280 1.0000 0.9304
-0.750 -0.3009 0.03776 0.02823 0.0308 1.0000 0.9361
-0.500 -0.2716 0.03739 0.02771 0.0284 1.0000 0.9415
-0.250 -0.2634 0.03687 0.02711 0.0298 1.0000 0.9468
0.000 -0.2409 0.03656 0.02668 0.0286 1.0000 0.9517
0.250 -0.2201 0.03630 0.02634 0.0276 1.0000 0.9565
0.500 -0.2063 0.03603 0.02601 0.0279 1.0000 0.9617
0.750 -0.1754 0.03606 0.02595 0.0249 1.0000 0.9662
1.000 -0.1532 0.03605 0.02589 0.0236 1.0000 0.9713
1.250 -0.1264 0.03617 0.02595 0.0212 1.0000 0.9760
1.500 -0.0986 0.03637 0.02613 0.0187 1.0000 0.9810
1.750 -0.0726 0.03662 0.02637 0.0164 1.0000 0.9863
2.000 -0.0463 0.03692 0.02666 0.0139 1.0000 0.9916
2.250 -0.0233 0.03726 0.02702 0.0119 1.0000 0.9979
2.500 -0.0172 0.03727 0.02706 0.0129 1.0000 1.0000
2.750 -0.0175 0.03708 0.02691 0.0151 1.0000 1.0000
3.000 -0.0174 0.03690 0.02676 0.0171 1.0000 1.0000
3.250 -0.0163 0.03677 0.02667 0.0189 1.0000 1.0000
3.500 -0.0108 0.03685 0.02679 0.0197 1.0000 1.0000
3.750 0.0004 0.03723 0.02721 0.0195 1.0000 1.0000
4.000 0.0158 0.03789 0.02792 0.0183 1.0000 1.0000
4.250 0.1214 0.04245 0.03255 0.0015 0.9491 1.0000
4.500 0.1744 0.04374 0.03392 -0.0050 0.9179 1.0000
4.750 0.2235 0.04489 0.03519 -0.0107 0.8941 1.0000
5.000 0.2623 0.04560 0.03601 -0.0146 0.8706 1.0000
5.250 0.3127 0.04651 0.03706 -0.0200 0.8493 1.0000
5.500 0.3523 0.04709 0.03782 -0.0236 0.8267 1.0000
5.750 0.4032 0.04768 0.03859 -0.0285 0.8062 1.0000
6.000 0.4436 0.04806 0.03916 -0.0317 0.7834 1.0000
6.250 0.4962 0.04808 0.03945 -0.0360 0.7620 1.0000
6.500 0.5411 0.04794 0.03955 -0.0389 0.7382 1.0000
6.750 0.5892 0.04740 0.03932 -0.0416 0.7140 1.0000
7.000 0.6577 0.04530 0.03762 -0.0452 0.6911 1.0000
7.250 0.7108 0.04334 0.03605 -0.0466 0.6641 1.0000
7.500 0.7919 0.03831 0.03154 -0.0484 0.6371 1.0000
7.750 0.9149 0.03030 0.02406 -0.0527 0.5820 1.0000
8.000 0.9603 0.02895 0.02254 -0.0521 0.5014 1.0000
8.250 0.9792 0.02956 0.02274 -0.0494 0.4212 1.0000
8.500 0.9869 0.03108 0.02374 -0.0461 0.3457 1.0000
8.750 0.9872 0.03302 0.02512 -0.0425 0.2776 1.0000
9.000 0.9834 0.03527 0.02686 -0.0390 0.2180 1.0000
9.250 0.9893 0.03817 0.02921 -0.0368 0.1602 1.0000
9.500 1.0314 0.04184 0.03251 -0.0389 0.1174 1.0000
9.750 1.0888 0.04606 0.03697 -0.0430 0.0999 1.0000
10.000 1.1297 0.05023 0.04131 -0.0458 0.0911 1.0000
10.250 1.1500 0.05425 0.04594 -0.0453 0.0881 1.0000
10.500 1.1623 0.05839 0.05059 -0.0441 0.0870 1.0000
10.750 1.1646 0.06247 0.05515 -0.0419 0.0866 1.0000
11.000 1.1590 0.06659 0.05969 -0.0392 0.0869 1.0000
11.250 1.1477 0.07077 0.06427 -0.0364 0.0875 1.0000
11.500 1.1319 0.07509 0.06893 -0.0338 0.0882 1.0000
11.750 1.1136 0.07964 0.07376 -0.0317 0.0890 1.0000
12.000 1.0922 0.08451 0.07889 -0.0304 0.0899 1.0000
12.250 1.0705 0.08982 0.08442 -0.0300 0.0908 1.0000
12.500 1.0519 0.09559 0.09034 -0.0305 0.0917 1.0000
12.750 1.0402 0.10174 0.09660 -0.0318 0.0926 1.0000
13.000 0.9264 0.11860 0.11378 -0.0449 0.0991 1.0000
13.250 0.9031 0.13032 0.12544 -0.0525 0.1034 1.0000
13.500 0.9067 0.13678 0.13192 -0.0550 0.1056 1.0000
13.750 0.8767 0.15179 0.14679 -0.0658 0.1185 1.0000
14.000 0.6112 0.15688 0.15236 -0.0579 0.2498 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 417 AIRFOIL (e417-il)