EPPLER 49 AIRFOIL (e49-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 49 AIRFOIL (e49-il) Reynolds number: 200,000 Max Cl/Cd: 65 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e49-il-200000-n5.txt Download as CSV file: xf-e49-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 49 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.2129 0.10853 0.10500 -0.0611 0.9323 0.0167
-7.750 -0.2161 0.10659 0.10309 -0.0599 0.9290 0.0167
-7.500 -0.2117 0.10395 0.10047 -0.0609 0.9262 0.0167
-7.250 -0.1968 0.10045 0.09688 -0.0646 0.9242 0.0167
-7.000 -0.1962 0.09781 0.09427 -0.0643 0.9181 0.0167
-6.750 -0.1870 0.09461 0.09108 -0.0663 0.9143 0.0167
-6.500 -0.1708 0.09089 0.08736 -0.0700 0.9119 0.0167
-6.250 -0.1493 0.08680 0.08326 -0.0749 0.9101 0.0167
-6.000 -0.1342 0.08328 0.07973 -0.0780 0.9069 0.0167
-5.750 -0.1263 0.08021 0.07667 -0.0790 0.9015 0.0167
-5.500 -0.1027 0.07608 0.07251 -0.0838 0.8986 0.0166
-5.250 -0.0738 0.07160 0.06798 -0.0897 0.8967 0.0165
-5.000 -0.0435 0.06701 0.06335 -0.0956 0.8954 0.0162
-4.750 -0.0091 0.06225 0.05853 -0.1024 0.8944 0.0155
-4.500 0.0140 0.05856 0.05477 -0.1062 0.8892 0.0151
-4.250 0.0577 0.05379 0.04987 -0.1145 0.8868 0.0150
-3.000 0.2978 0.03489 0.02958 -0.1488 0.8821 0.0226
-2.750 0.3440 0.03191 0.02604 -0.1528 0.8815 0.0194
-2.250 0.4206 0.02934 0.02254 -0.1567 0.8795 0.0173
-2.000 0.4580 0.02785 0.02068 -0.1589 0.8788 0.0174
-1.750 0.4760 0.02675 0.01932 -0.1573 0.8720 0.0182
-1.500 0.5058 0.02590 0.01831 -0.1579 0.8690 0.0199
-1.250 0.5376 0.02531 0.01758 -0.1589 0.8667 0.0244
-1.000 0.5702 0.02479 0.01696 -0.1600 0.8649 0.0301
-0.750 0.6040 0.02421 0.01635 -0.1612 0.8635 0.0293
-0.500 0.6223 0.02403 0.01616 -0.1596 0.8562 0.0289
-0.250 0.6505 0.02371 0.01582 -0.1599 0.8520 0.0285
0.000 0.6828 0.02336 0.01545 -0.1609 0.8495 0.0290
0.250 0.7165 0.02301 0.01503 -0.1621 0.8476 0.0295
0.750 0.7655 0.02266 0.01463 -0.1610 0.8346 0.0357
1.000 0.7992 0.02072 0.01465 -0.1626 0.8323 1.0000
1.250 0.8193 0.02067 0.01453 -0.1611 0.8215 1.0000
1.500 0.8518 0.02022 0.01403 -0.1618 0.8167 1.0000
1.750 0.8713 0.02029 0.01409 -0.1603 0.8056 1.0000
2.000 0.9013 0.01997 0.01375 -0.1607 0.8008 1.0000
2.250 0.9218 0.02007 0.01386 -0.1594 0.7894 1.0000
2.500 0.9528 0.01971 0.01358 -0.1600 0.7836 1.0000
3.000 1.0083 0.01926 0.01314 -0.1599 0.7624 1.0000
3.250 1.0462 0.01849 0.01235 -0.1616 0.7494 1.0000
3.500 1.0931 0.01739 0.01121 -0.1649 0.7316 1.0000
3.750 1.1355 0.01747 0.00905 -0.1666 0.4209 1.0000
4.000 1.1226 0.02000 0.01041 -0.1598 0.2439 1.0000
4.250 1.1406 0.02064 0.01103 -0.1584 0.2148 1.0000
4.500 1.1508 0.02203 0.01192 -0.1559 0.1332 1.0000
4.750 1.1547 0.02417 0.01334 -0.1524 0.0071 1.0000
5.000 1.1745 0.02472 0.01397 -0.1512 0.0063 1.0000
5.250 1.1938 0.02533 0.01468 -0.1499 0.0058 1.0000
5.500 1.2125 0.02600 0.01548 -0.1486 0.0056 1.0000
5.750 1.2304 0.02676 0.01639 -0.1471 0.0054 1.0000
6.000 1.2475 0.02758 0.01736 -0.1455 0.0054 1.0000
6.250 1.2637 0.02852 0.01847 -0.1437 0.0053 1.0000
6.500 1.2785 0.02962 0.01983 -0.1418 0.0052 1.0000
6.750 1.2919 0.03083 0.02120 -0.1397 0.0052 1.0000
7.000 1.3035 0.03219 0.02271 -0.1373 0.0053 1.0000
7.250 1.3132 0.03370 0.02437 -0.1348 0.0053 1.0000
7.500 1.3223 0.03532 0.02612 -0.1322 0.0054 1.0000
7.750 1.3316 0.03707 0.02800 -0.1297 0.0055 1.0000
8.000 1.3421 0.03891 0.02996 -0.1275 0.0056 1.0000
8.250 1.3549 0.04079 0.03197 -0.1256 0.0057 1.0000
8.500 1.3709 0.04275 0.03407 -0.1242 0.0058 1.0000
8.750 1.3903 0.04483 0.03632 -0.1232 0.0059 1.0000
9.000 1.4115 0.04712 0.03882 -0.1225 0.0061 1.0000
9.250 1.4315 0.04968 0.04164 -0.1216 0.0062 1.0000
9.500 1.4474 0.05255 0.04481 -0.1203 0.0063 1.0000
9.750 1.4575 0.05566 0.04823 -0.1184 0.0065 1.0000
10.000 1.4621 0.05900 0.05191 -0.1159 0.0066 1.0000
10.250 1.4614 0.06254 0.05578 -0.1131 0.0067 1.0000
10.500 1.4566 0.06628 0.05982 -0.1101 0.0068 1.0000
10.750 1.4483 0.07025 0.06410 -0.1071 0.0069 1.0000
11.000 1.4372 0.07443 0.06858 -0.1043 0.0070 1.0000
11.250 1.4238 0.07883 0.07326 -0.1019 0.0070 1.0000
11.500 1.4088 0.08349 0.07818 -0.0999 0.0071 1.0000
11.750 1.3926 0.08839 0.08334 -0.0986 0.0071 1.0000
12.000 1.3757 0.09358 0.08877 -0.0979 0.0072 1.0000
12.250 1.3583 0.09911 0.09453 -0.0980 0.0072 1.0000
12.500 1.3410 0.10490 0.10053 -0.0989 0.0073 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 49 AIRFOIL (e49-il)