EPPLER 421 AIRFOIL (e421-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 421 AIRFOIL (e421-il) Reynolds number: 500,000 Max Cl/Cd: 107.39 at α=6.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e421-il-500000.txt Download as CSV file: xf-e421-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 421 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 0.2320 0.09015 0.08604 -0.1303 0.6594 0.0236
-9.250 0.2394 0.08812 0.08394 -0.1306 0.6496 0.0242
-9.000 0.2472 0.08599 0.08180 -0.1312 0.6408 0.0244
-8.750 0.2526 0.08357 0.07934 -0.1318 0.6326 0.0253
-8.500 0.2426 0.07938 0.07516 -0.1345 0.6276 0.0261
-8.250 0.2542 0.07757 0.07334 -0.1343 0.6196 0.0263
-8.000 0.2644 0.07586 0.07157 -0.1341 0.6123 0.0265
-7.750 0.2756 0.07427 0.06997 -0.1340 0.6059 0.0268
-7.500 0.2838 0.07240 0.06810 -0.1342 0.5995 0.0273
-7.250 0.1193 0.02393 0.01882 -0.1777 0.6113 0.0155
-7.000 0.1433 0.02109 0.01554 -0.1803 0.6049 0.0158
-6.750 0.1679 0.01952 0.01363 -0.1811 0.5981 0.0160
-6.500 0.1921 0.01813 0.01191 -0.1815 0.5918 0.0164
-6.250 0.2177 0.01721 0.01091 -0.1817 0.5862 0.0169
-6.000 0.2437 0.01656 0.01013 -0.1816 0.5802 0.0174
-5.750 0.2698 0.01602 0.00942 -0.1815 0.5745 0.0180
-5.500 0.2965 0.01544 0.00870 -0.1814 0.5694 0.0186
-5.250 0.3236 0.01496 0.00808 -0.1813 0.5641 0.0192
-5.000 0.3499 0.01436 0.00740 -0.1812 0.5590 0.0201
-4.750 0.3768 0.01406 0.00698 -0.1810 0.5540 0.0212
-4.500 0.4045 0.01367 0.00651 -0.1809 0.5501 0.0226
-4.250 0.4318 0.01334 0.00615 -0.1808 0.5458 0.0249
-4.000 0.4592 0.01303 0.00577 -0.1807 0.5417 0.0280
-3.750 0.4866 0.01280 0.00545 -0.1806 0.5376 0.0328
-3.500 0.5145 0.01254 0.00515 -0.1805 0.5339 0.0402
-3.250 0.5421 0.01234 0.00494 -0.1803 0.5303 0.0499
-3.000 0.5696 0.01218 0.00480 -0.1802 0.5265 0.0597
-2.750 0.5971 0.01211 0.00470 -0.1800 0.5230 0.0692
-2.500 0.6247 0.01210 0.00463 -0.1799 0.5196 0.0788
-2.250 0.6526 0.01212 0.00462 -0.1798 0.5164 0.0890
-2.000 0.6802 0.01206 0.00459 -0.1796 0.5137 0.0995
-1.750 0.7077 0.01203 0.00458 -0.1795 0.5108 0.1102
-1.500 0.7352 0.01201 0.00454 -0.1792 0.5077 0.1201
-1.250 0.7625 0.01207 0.00456 -0.1790 0.5047 0.1299
-1.000 0.7900 0.01212 0.00458 -0.1789 0.5017 0.1405
-0.750 0.8182 0.01222 0.00464 -0.1789 0.4988 0.1501
-0.500 0.8454 0.01224 0.00469 -0.1786 0.4967 0.1596
-0.250 0.8726 0.01224 0.00471 -0.1784 0.4943 0.1690
0.000 0.8998 0.01231 0.00476 -0.1782 0.4917 0.1782
0.250 0.9269 0.01233 0.00478 -0.1780 0.4891 0.1876
0.500 0.9539 0.01241 0.00484 -0.1777 0.4865 0.1967
0.750 0.9814 0.01250 0.00490 -0.1776 0.4840 0.2060
1.000 1.0098 0.01270 0.00505 -0.1777 0.4812 0.2156
1.250 1.0363 0.01271 0.00511 -0.1774 0.4796 0.2251
1.500 1.0629 0.01276 0.00519 -0.1772 0.4775 0.2350
1.750 1.0893 0.01280 0.00526 -0.1769 0.4751 0.2454
2.000 1.1159 0.01288 0.00535 -0.1766 0.4727 0.2563
2.250 1.1424 0.01293 0.00543 -0.1764 0.4705 0.2684
2.500 1.1691 0.01303 0.00554 -0.1762 0.4684 0.2823
2.750 1.1965 0.01318 0.00569 -0.1761 0.4662 0.2982
3.000 1.2250 0.01340 0.00592 -0.1764 0.4637 0.3189
3.250 1.2502 0.01342 0.00606 -0.1759 0.4621 0.3456
3.500 1.2755 0.01344 0.00621 -0.1755 0.4600 0.3853
3.750 1.3008 0.01331 0.00642 -0.1753 0.4579 0.5212
4.000 1.3247 0.01276 0.00661 -0.1745 0.4558 1.0000
4.250 1.3506 0.01292 0.00674 -0.1742 0.4538 1.0000
4.500 1.3764 0.01310 0.00688 -0.1739 0.4519 1.0000
4.750 1.4027 0.01331 0.00704 -0.1737 0.4497 1.0000
5.000 1.4310 0.01365 0.00730 -0.1739 0.4471 1.0000
5.250 1.4554 0.01382 0.00750 -0.1734 0.4454 1.0000
5.500 1.4792 0.01397 0.00769 -0.1727 0.4436 1.0000
5.750 1.5033 0.01415 0.00788 -0.1721 0.4416 1.0000
6.000 1.5273 0.01432 0.00806 -0.1715 0.4394 1.0000
6.250 1.5514 0.01449 0.00825 -0.1710 0.4372 1.0000
6.500 1.5755 0.01468 0.00842 -0.1704 0.4351 1.0000
6.750 1.6001 0.01490 0.00862 -0.1700 0.4330 1.0000
7.000 1.6278 0.01525 0.00893 -0.1703 0.4303 1.0000
7.250 1.6497 0.01546 0.00919 -0.1694 0.4284 1.0000
7.500 1.6695 0.01562 0.00941 -0.1680 0.4261 1.0000
7.750 1.6894 0.01579 0.00963 -0.1667 0.4236 1.0000
8.000 1.7087 0.01595 0.00981 -0.1653 0.4210 1.0000
8.250 1.7283 0.01613 0.01000 -0.1640 0.4185 1.0000
8.500 1.7494 0.01636 0.01022 -0.1629 0.4158 1.0000
8.750 1.7749 0.01672 0.01055 -0.1629 0.4128 1.0000
9.000 1.7883 0.01691 0.01084 -0.1604 0.4107 1.0000
9.250 1.8037 0.01713 0.01114 -0.1584 0.4079 1.0000
9.500 1.8202 0.01736 0.01142 -0.1566 0.4051 1.0000
9.750 1.8374 0.01761 0.01169 -0.1551 0.4024 1.0000
10.000 1.8559 0.01789 0.01198 -0.1537 0.3996 1.0000
10.250 1.8763 0.01826 0.01234 -0.1529 0.3965 1.0000
10.500 1.8873 0.01859 0.01279 -0.1503 0.3937 1.0000
10.750 1.9002 0.01895 0.01323 -0.1481 0.3904 1.0000
11.000 1.9134 0.01932 0.01365 -0.1461 0.3870 1.0000
11.250 1.9280 0.01973 0.01406 -0.1444 0.3837 1.0000
11.500 1.9431 0.02022 0.01458 -0.1428 0.3803 1.0000
11.750 1.9527 0.02078 0.01526 -0.1405 0.3770 1.0000
12.000 1.9631 0.02137 0.01593 -0.1384 0.3732 1.0000
12.250 1.9737 0.02201 0.01660 -0.1364 0.3694 1.0000
12.500 1.9852 0.02272 0.01732 -0.1346 0.3654 1.0000
12.750 1.9925 0.02359 0.01832 -0.1324 0.3613 1.0000
13.000 1.9999 0.02452 0.01932 -0.1304 0.3567 1.0000
13.250 2.0064 0.02557 0.02038 -0.1283 0.3521 1.0000
13.500 2.0122 0.02676 0.02167 -0.1264 0.3472 1.0000
13.750 2.0164 0.02810 0.02309 -0.1245 0.3415 1.0000
14.000 2.0184 0.02967 0.02465 -0.1225 0.3363 1.0000
14.250 2.0221 0.03127 0.02639 -0.1209 0.3304 1.0000
14.500 2.0211 0.03327 0.02842 -0.1190 0.3240 1.0000
14.750 2.0201 0.03541 0.03063 -0.1174 0.3174 1.0000
15.000 2.0153 0.03798 0.03324 -0.1157 0.3099 1.0000
15.250 2.0106 0.04068 0.03601 -0.1143 0.3025 1.0000
15.500 2.0009 0.04394 0.03929 -0.1128 0.2944 1.0000
15.750 1.9919 0.04732 0.04274 -0.1116 0.2859 1.0000
16.000 1.9762 0.05149 0.04692 -0.1104 0.2772 1.0000
16.250 1.9633 0.05556 0.05105 -0.1095 0.2679 1.0000
16.500 1.9475 0.06006 0.05558 -0.1087 0.2589 1.0000
16.750 1.9269 0.06527 0.06079 -0.1081 0.2493 1.0000
17.000 1.9110 0.07007 0.06564 -0.1077 0.2398 1.0000
17.250 1.8940 0.07511 0.07071 -0.1076 0.2310 1.0000
17.500 1.8745 0.08059 0.07618 -0.1076 0.2222 1.0000
17.750 1.8624 0.08524 0.08089 -0.1077 0.2138 1.0000
18.000 1.8474 0.09030 0.08595 -0.1080 0.2057 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 421 AIRFOIL (e421-il)