EPPLER 417 AIRFOIL (e417-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 417 AIRFOIL (e417-il) Reynolds number: 50,000 Max Cl/Cd: 30.21 at α=7.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e417-il-50000-n5.txt Download as CSV file: xf-e417-il-50000-n5.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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.5277 0.10765 0.10068 -0.0482 1.0000 0.0425
-10.500 -0.5471 0.10038 0.09346 -0.0512 1.0000 0.0416
-10.250 -0.5728 0.09341 0.08650 -0.0540 1.0000 0.0407
-10.000 -0.6041 0.08719 0.08025 -0.0559 1.0000 0.0400
-9.750 -0.6382 0.08210 0.07507 -0.0563 1.0000 0.0394
-9.500 -0.6752 0.07829 0.07113 -0.0550 1.0000 0.0388
-9.250 -0.7070 0.07536 0.06809 -0.0525 1.0000 0.0385
-9.000 -0.7270 0.07207 0.06459 -0.0511 1.0000 0.0382
-8.750 -0.7395 0.06879 0.06106 -0.0496 1.0000 0.0380
-8.500 -0.7457 0.06545 0.05740 -0.0483 1.0000 0.0378
-8.250 -0.7468 0.06207 0.05375 -0.0472 1.0000 0.0377
-8.000 -0.7446 0.05867 0.05002 -0.0462 1.0000 0.0378
-7.750 -0.7379 0.05549 0.04650 -0.0454 1.0000 0.0377
-7.250 -0.7147 0.04952 0.03988 -0.0438 1.0000 0.0381
-7.000 -0.6993 0.04690 0.03700 -0.0429 1.0000 0.0385
-6.750 -0.6827 0.04469 0.03458 -0.0420 1.0000 0.0395
-6.500 -0.6652 0.04292 0.03264 -0.0410 1.0000 0.0412
-6.250 -0.6467 0.04136 0.03090 -0.0398 1.0000 0.0430
-6.000 -0.6276 0.03999 0.02933 -0.0381 1.0000 0.0449
-5.750 -0.6086 0.03887 0.02802 -0.0358 1.0000 0.0465
-5.500 -0.5903 0.03801 0.02697 -0.0332 1.0000 0.0478
-5.250 -0.5741 0.03706 0.02595 -0.0306 1.0000 0.0496
-5.000 -0.5582 0.03605 0.02492 -0.0290 1.0000 0.0521
-4.750 -0.5410 0.03505 0.02382 -0.0278 1.0000 0.0559
-4.500 -0.5224 0.03393 0.02262 -0.0273 1.0000 0.0626
-4.250 -0.5016 0.03270 0.02135 -0.0276 1.0000 0.0737
-4.000 -0.4762 0.03066 0.01965 -0.0297 1.0000 0.1089
-3.250 -0.4574 0.03387 0.02570 -0.0114 1.0000 0.6492
-3.000 -0.4452 0.03511 0.02675 -0.0063 0.9978 0.6756
-2.750 -0.4217 0.03581 0.02716 -0.0046 0.9930 0.6923
-2.500 -0.3906 0.03588 0.02691 -0.0061 0.9883 0.7037
-2.250 -0.3631 0.03602 0.02679 -0.0063 0.9836 0.7104
-2.000 -0.3301 0.03588 0.02635 -0.0089 0.9787 0.7186
-1.750 -0.3000 0.03600 0.02624 -0.0099 0.9740 0.7236
-1.500 -0.2707 0.03584 0.02587 -0.0114 0.9683 0.7294
-1.250 -0.2362 0.03587 0.02568 -0.0141 0.9634 0.7349
-1.000 -0.2090 0.03579 0.02544 -0.0148 0.9574 0.7391
-0.750 -0.1758 0.03582 0.02530 -0.0171 0.9519 0.7442
-0.500 -0.1423 0.03580 0.02513 -0.0196 0.9462 0.7489
-0.250 -0.1138 0.03581 0.02503 -0.0205 0.9397 0.7526
0.000 -0.0820 0.03588 0.02499 -0.0223 0.9338 0.7568
0.250 -0.0500 0.03588 0.02489 -0.0245 0.9268 0.7613
0.500 -0.0153 0.03601 0.02493 -0.0268 0.9210 0.7652
0.750 0.0107 0.03597 0.02486 -0.0273 0.9130 0.7688
1.000 0.0460 0.03611 0.02494 -0.0298 0.9068 0.7727
1.250 0.0772 0.03613 0.02493 -0.0317 0.8985 0.7770
1.500 0.1096 0.03621 0.02499 -0.0334 0.8913 0.7805
1.750 0.1397 0.03624 0.02504 -0.0346 0.8830 0.7840
2.000 0.1698 0.03628 0.02509 -0.0359 0.8743 0.7879
2.250 0.2085 0.03637 0.02520 -0.0389 0.8672 0.7920
2.500 0.2349 0.03636 0.02526 -0.0394 0.8569 0.7955
2.750 0.2662 0.03635 0.02530 -0.0407 0.8480 0.7991
3.000 0.3011 0.03631 0.02534 -0.0426 0.8391 0.8030
3.250 0.3313 0.03633 0.02545 -0.0439 0.8279 0.8070
3.500 0.3607 0.03625 0.02548 -0.0447 0.8173 0.8107
3.750 0.3976 0.03605 0.02540 -0.0465 0.8090 0.8145
4.000 0.4253 0.03602 0.02550 -0.0471 0.7965 0.8186
4.250 0.4563 0.03599 0.02563 -0.0483 0.7840 0.8228
4.500 0.4840 0.03579 0.02559 -0.0484 0.7714 0.8266
4.750 0.5154 0.03551 0.02548 -0.0490 0.7586 0.8308
5.000 0.5500 0.03513 0.02531 -0.0502 0.7454 0.8352
5.250 0.5841 0.03461 0.02500 -0.0509 0.7321 0.8394
5.500 0.6177 0.03392 0.02454 -0.0512 0.7189 0.8439
5.750 0.6478 0.03344 0.02430 -0.0512 0.7023 0.8487
6.000 0.6781 0.03293 0.02403 -0.0512 0.6842 0.8534
6.250 0.7090 0.03220 0.02355 -0.0507 0.6651 0.8582
6.500 0.7470 0.03120 0.02282 -0.0511 0.6446 0.8634
6.750 0.7775 0.03063 0.02246 -0.0507 0.6173 0.8687
7.000 0.8111 0.02987 0.02188 -0.0504 0.5855 0.8743
7.250 0.8444 0.02941 0.02154 -0.0503 0.5458 0.8802
7.500 0.8709 0.02923 0.02137 -0.0492 0.5009 0.8861
7.750 0.8920 0.02953 0.02161 -0.0477 0.4521 0.8930
8.000 0.9072 0.03013 0.02212 -0.0457 0.4039 0.9002
8.250 0.9187 0.03103 0.02291 -0.0434 0.3550 0.9084
8.500 0.9263 0.03210 0.02385 -0.0408 0.3078 0.9176
8.750 0.9325 0.03339 0.02499 -0.0385 0.2620 0.9287
9.000 0.9373 0.03493 0.02635 -0.0363 0.2158 0.9438
9.500 0.9453 0.03901 0.02981 -0.0337 0.1168 1.0000
9.750 0.9542 0.04159 0.03214 -0.0333 0.0861 1.0000
10.000 0.9647 0.04415 0.03470 -0.0328 0.0674 1.0000
10.250 0.9768 0.04665 0.03725 -0.0324 0.0578 1.0000
10.500 0.9899 0.04909 0.03973 -0.0321 0.0523 1.0000
10.750 1.0107 0.05129 0.04218 -0.0320 0.0477 1.0000
11.000 1.0281 0.05359 0.04463 -0.0320 0.0437 1.0000
11.250 1.0455 0.05607 0.04716 -0.0321 0.0409 1.0000
11.500 1.0672 0.05873 0.05025 -0.0321 0.0384 1.0000
11.750 1.0848 0.06179 0.05365 -0.0321 0.0370 1.0000
12.000 1.0959 0.06518 0.05737 -0.0318 0.0360 1.0000
12.250 1.1009 0.06888 0.06139 -0.0315 0.0353 1.0000
12.500 1.1005 0.07286 0.06569 -0.0312 0.0348 1.0000
12.750 1.0951 0.07720 0.07034 -0.0310 0.0345 1.0000
13.000 1.0853 0.08198 0.07542 -0.0313 0.0343 1.0000
13.250 1.0712 0.08731 0.08105 -0.0323 0.0343 1.0000
13.500 1.0533 0.09336 0.08740 -0.0342 0.0344 1.0000
13.750 1.0315 0.10040 0.09472 -0.0373 0.0346 1.0000
14.000 1.0041 0.10916 0.10375 -0.0424 0.0350 1.0000
14.250 0.9723 0.12028 0.11510 -0.0499 0.0358 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 417 AIRFOIL (e417-il)