EPPLER 397 AIRFOIL (e397-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 397 AIRFOIL (e397-il) Reynolds number: 200,000 Max Cl/Cd: 89.05 at α=7.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e397-il-200000.txt Download as CSV file: xf-e397-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 397 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.500 -0.2160 0.10500 0.10167 -0.0837 0.9728 0.0553
-10.250 -0.3601 0.05808 0.05450 -0.1146 0.9600 0.0239
-10.000 -0.3717 0.04952 0.04572 -0.1267 0.9488 0.0237
-9.750 -0.3765 0.04233 0.03816 -0.1386 0.9368 0.0236
-9.500 -0.3637 0.03626 0.03143 -0.1477 0.9307 0.0237
-9.250 -0.3540 0.03275 0.02743 -0.1498 0.9219 0.0241
-9.000 -0.3290 0.02968 0.02419 -0.1525 0.9178 0.0253
-8.750 -0.2969 0.02746 0.02171 -0.1556 0.9152 0.0272
-8.500 -0.2784 0.02602 0.01989 -0.1554 0.9071 0.0294
-8.250 -0.2494 0.02441 0.01828 -0.1572 0.9031 0.0333
-8.000 -0.2172 0.02259 0.01622 -0.1593 0.9002 0.0384
-7.750 -0.1958 0.02198 0.01545 -0.1588 0.8931 0.0441
-7.500 -0.1673 0.02092 0.01435 -0.1599 0.8885 0.0520
-7.250 -0.1346 0.01994 0.01329 -0.1615 0.8853 0.0615
-7.000 -0.1117 0.01928 0.01256 -0.1612 0.8792 0.0703
-6.750 -0.0843 0.01860 0.01177 -0.1616 0.8743 0.0805
-6.500 -0.0526 0.01797 0.01102 -0.1627 0.8709 0.0922
-6.250 -0.0247 0.01759 0.01055 -0.1631 0.8663 0.1043
-6.000 -0.0002 0.01715 0.01010 -0.1629 0.8605 0.1174
-5.750 0.0302 0.01669 0.00958 -0.1636 0.8568 0.1330
-5.500 0.0623 0.01616 0.00907 -0.1648 0.8540 0.1506
-5.250 0.0842 0.01601 0.00894 -0.1640 0.8477 0.1666
-5.000 0.1128 0.01575 0.00867 -0.1644 0.8433 0.1857
-4.750 0.1442 0.01551 0.00836 -0.1652 0.8402 0.2069
-4.500 0.1717 0.01536 0.00823 -0.1654 0.8361 0.2267
-4.250 0.1966 0.01530 0.00818 -0.1651 0.8308 0.2458
-4.000 0.2263 0.01516 0.00801 -0.1656 0.8270 0.2672
-3.750 0.2578 0.01501 0.00783 -0.1663 0.8242 0.2887
-3.500 0.2824 0.01504 0.00790 -0.1660 0.8194 0.3076
-3.250 0.3089 0.01503 0.00789 -0.1659 0.8147 0.3275
-3.000 0.3391 0.01496 0.00776 -0.1663 0.8112 0.3488
-2.750 0.3703 0.01484 0.00766 -0.1670 0.8086 0.3691
-2.500 0.3934 0.01495 0.00783 -0.1664 0.8034 0.3877
-2.250 0.4204 0.01497 0.00786 -0.1663 0.7990 0.4079
-2.000 0.4505 0.01491 0.00779 -0.1668 0.7957 0.4296
-1.750 0.4818 0.01483 0.00774 -0.1674 0.7931 0.4509
-1.500 0.5039 0.01501 0.00799 -0.1665 0.7877 0.4714
-1.250 0.5308 0.01505 0.00806 -0.1665 0.7835 0.4937
-1.000 0.5607 0.01499 0.00806 -0.1668 0.7803 0.5171
-0.750 0.5922 0.01494 0.00802 -0.1674 0.7778 0.5425
-0.500 0.6129 0.01517 0.00837 -0.1663 0.7721 0.5658
-0.250 0.6395 0.01521 0.00850 -0.1661 0.7680 0.5920
0.000 0.6694 0.01517 0.00850 -0.1664 0.7648 0.6208
0.250 0.7004 0.01512 0.00850 -0.1669 0.7623 0.6514
0.500 0.7187 0.01541 0.00895 -0.1653 0.7563 0.6799
0.750 0.7446 0.01543 0.00907 -0.1648 0.7521 0.7124
1.000 0.7728 0.01536 0.00908 -0.1646 0.7491 0.7465
1.250 0.7967 0.01541 0.00924 -0.1636 0.7454 0.7823
1.500 0.8128 0.01560 0.00958 -0.1612 0.7395 0.8221
1.750 0.8326 0.01549 0.00957 -0.1591 0.7356 0.8700
2.000 0.8582 0.01519 0.00934 -0.1580 0.7327 0.9595
2.250 0.8849 0.01553 0.00968 -0.1584 0.7269 1.0000
2.500 0.9143 0.01567 0.00978 -0.1590 0.7220 1.0000
2.750 0.9485 0.01564 0.00969 -0.1603 0.7186 1.0000
3.000 0.9755 0.01586 0.00991 -0.1603 0.7136 1.0000
3.250 1.0006 0.01607 0.01013 -0.1600 0.7076 1.0000
3.500 1.0339 0.01600 0.01004 -0.1609 0.7037 1.0000
3.750 1.0604 0.01617 0.01021 -0.1608 0.6982 1.0000
4.000 1.0860 0.01629 0.01036 -0.1605 0.6918 1.0000
4.250 1.1202 0.01615 0.01019 -0.1615 0.6876 1.0000
4.500 1.1418 0.01640 0.01052 -0.1604 0.6804 1.0000
4.750 1.1719 0.01632 0.01044 -0.1607 0.6745 1.0000
5.000 1.2002 0.01632 0.01047 -0.1607 0.6682 1.0000
5.250 1.2258 0.01628 0.01047 -0.1602 0.6601 1.0000
5.500 1.2529 0.01616 0.01037 -0.1598 0.6517 1.0000
5.750 1.2826 0.01584 0.01003 -0.1597 0.6421 1.0000
6.000 1.3031 0.01577 0.01003 -0.1581 0.6302 1.0000
6.250 1.3258 0.01570 0.00999 -0.1570 0.6190 1.0000
6.500 1.3509 0.01563 0.00994 -0.1562 0.6085 1.0000
6.750 1.3728 0.01569 0.01007 -0.1550 0.5979 1.0000
7.000 1.3923 0.01584 0.01031 -0.1535 0.5867 1.0000
7.250 1.4130 0.01596 0.01049 -0.1521 0.5749 1.0000
7.500 1.4328 0.01609 0.01069 -0.1505 0.5619 1.0000
7.750 1.4495 0.01628 0.01091 -0.1484 0.5462 1.0000
8.000 1.4635 0.01652 0.01118 -0.1458 0.5280 1.0000
8.250 1.4746 0.01682 0.01146 -0.1426 0.5082 1.0000
8.500 1.4815 0.01727 0.01189 -0.1388 0.4856 1.0000
8.750 1.4873 0.01786 0.01241 -0.1350 0.4610 1.0000
9.000 1.4912 0.01862 0.01311 -0.1311 0.4349 1.0000
9.250 1.4924 0.01957 0.01397 -0.1269 0.4070 1.0000
9.500 1.4895 0.02082 0.01507 -0.1225 0.3762 1.0000
9.750 1.4846 0.02232 0.01643 -0.1182 0.3451 1.0000
10.000 1.4792 0.02403 0.01800 -0.1141 0.3150 1.0000
10.250 1.4737 0.02594 0.01976 -0.1104 0.2855 1.0000
10.500 1.4689 0.02797 0.02167 -0.1071 0.2577 1.0000
10.750 1.4647 0.03013 0.02371 -0.1042 0.2313 1.0000
11.000 1.4602 0.03244 0.02592 -0.1015 0.2063 1.0000
11.250 1.4568 0.03482 0.02820 -0.0991 0.1805 1.0000
11.500 1.4526 0.03740 0.03066 -0.0969 0.1550 1.0000
11.750 1.4474 0.04018 0.03332 -0.0948 0.1304 1.0000
12.000 1.4418 0.04316 0.03616 -0.0929 0.1096 1.0000
12.250 1.4381 0.04608 0.03900 -0.0913 0.0940 1.0000
12.500 1.4357 0.04899 0.04186 -0.0899 0.0823 1.0000
12.750 1.4371 0.05161 0.04449 -0.0888 0.0722 1.0000
13.000 1.4384 0.05430 0.04721 -0.0879 0.0639 1.0000
13.250 1.4361 0.05748 0.05039 -0.0869 0.0574 1.0000
13.500 1.4326 0.06089 0.05382 -0.0861 0.0518 1.0000
13.750 1.4293 0.06438 0.05739 -0.0853 0.0468 1.0000
14.000 1.4258 0.06796 0.06101 -0.0847 0.0425 1.0000
14.250 1.4230 0.07155 0.06468 -0.0841 0.0383 1.0000
14.500 1.4240 0.07478 0.06800 -0.0839 0.0345 1.0000
14.750 1.4219 0.07835 0.07157 -0.0834 0.0315 1.0000
15.000 1.4250 0.08146 0.07485 -0.0835 0.0285 1.0000
15.250 1.4254 0.08488 0.07829 -0.0837 0.0263 1.0000
15.500 1.4269 0.08823 0.08176 -0.0837 0.0241 1.0000
15.750 1.4289 0.09158 0.08522 -0.0840 0.0223 1.0000
16.000 1.4305 0.09492 0.08861 -0.0844 0.0210 1.0000
16.250 1.4339 0.09794 0.09169 -0.0842 0.0197 1.0000
16.500 1.4370 0.10121 0.09514 -0.0844 0.0186 1.0000
16.750 1.4385 0.10471 0.09877 -0.0851 0.0176 1.0000
17.000 1.4398 0.10819 0.10230 -0.0859 0.0168 1.0000
17.250 1.4429 0.11126 0.10540 -0.0861 0.0158 1.0000
17.500 1.4405 0.11559 0.10999 -0.0874 0.0152 1.0000
17.750 1.4386 0.11983 0.11443 -0.0888 0.0146 1.0000
18.000 1.4375 0.12393 0.11871 -0.0901 0.0142 1.0000
18.250 1.4352 0.12826 0.12322 -0.0917 0.0139 1.0000
18.500 1.4318 0.13286 0.12799 -0.0937 0.0135 1.0000
18.750 1.4280 0.13754 0.13283 -0.0958 0.0133 1.0000
19.000 1.4238 0.14232 0.13777 -0.0982 0.0131 1.0000
19.250 1.4188 0.14730 0.14289 -0.1008 0.0129 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 397 AIRFOIL (e397-il)