EPPLER 432 AIRFOIL (e432-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 432 AIRFOIL (e432-il) Reynolds number: 500,000 Max Cl/Cd: 112.21 at α=8° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e432-il-500000.txt Download as CSV file: xf-e432-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 432 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
------ -------- --------- --------- -------- -------- --------
-11.750 -0.2063 0.11785 0.11560 -0.0643 0.9905 0.0161
-11.500 -0.1949 0.11267 0.11043 -0.0690 0.9853 0.0170
-11.250 -0.1965 0.10476 0.10252 -0.0757 0.9735 0.0176
-10.750 -0.1639 0.09654 0.09428 -0.0836 0.9458 0.0182
-10.500 -0.1346 0.09169 0.08938 -0.0916 0.9350 0.0190
-10.250 -0.0236 0.06967 0.06682 -0.1100 0.8167 0.0222
-10.000 -0.0226 0.06640 0.06350 -0.1108 0.8027 0.0225
-9.750 -0.0243 0.06253 0.05958 -0.1120 0.7915 0.0228
-9.500 -0.0288 0.05819 0.05519 -0.1135 0.7819 0.0233
-9.250 -0.0370 0.05300 0.04998 -0.1160 0.7735 0.0237
-9.000 -0.0584 0.04498 0.04192 -0.1220 0.7672 0.0238
-8.750 -0.0899 0.03840 0.03525 -0.1258 0.7600 0.0232
-8.500 -0.1163 0.03442 0.03112 -0.1258 0.7522 0.0230
-8.250 -0.1386 0.03172 0.02830 -0.1234 0.7447 0.0230
-8.000 -0.1462 0.02912 0.02553 -0.1217 0.7379 0.0239
-7.750 -0.1775 0.02530 0.02092 -0.1171 0.7316 0.0257
-7.500 -0.1797 0.03405 0.02941 -0.1225 0.7372 0.0262
-7.250 -0.1881 0.02586 0.02025 -0.1180 0.7319 0.0179
-7.000 -0.1670 0.02428 0.01859 -0.1173 0.7258 0.0171
-6.750 -0.1526 0.02042 0.01398 -0.1151 0.7207 0.0145
-6.500 -0.1291 0.01912 0.01251 -0.1145 0.7158 0.0142
-6.250 -0.1060 0.01769 0.01094 -0.1138 0.7106 0.0139
-6.000 -0.0821 0.01660 0.00970 -0.1132 0.7056 0.0138
-5.750 -0.0578 0.01573 0.00868 -0.1127 0.7011 0.0138
-5.500 -0.0336 0.01498 0.00786 -0.1121 0.6967 0.0139
-5.250 -0.0093 0.01435 0.00715 -0.1115 0.6922 0.0141
-5.000 0.0143 0.01368 0.00642 -0.1110 0.6879 0.0145
-4.750 0.0386 0.01314 0.00582 -0.1106 0.6837 0.0153
-4.500 0.0637 0.01272 0.00540 -0.1103 0.6797 0.0165
-4.250 0.0891 0.01229 0.00493 -0.1100 0.6757 0.0181
-4.000 0.1150 0.01190 0.00453 -0.1098 0.6718 0.0227
-3.750 0.1414 0.01155 0.00416 -0.1097 0.6679 0.0349
-3.500 0.1671 0.01112 0.00391 -0.1096 0.6642 0.0668
-3.250 0.1934 0.01075 0.00373 -0.1097 0.6605 0.1146
-3.000 0.2194 0.01024 0.00354 -0.1099 0.6568 0.2013
-2.750 0.2450 0.00940 0.00330 -0.1105 0.6533 0.3837
-2.500 0.2700 0.00852 0.00334 -0.1108 0.6498 0.6490
-2.250 0.2977 0.00859 0.00345 -0.1106 0.6462 0.6958
-2.000 0.3257 0.00870 0.00354 -0.1105 0.6425 0.7164
-1.750 0.3538 0.00883 0.00361 -0.1104 0.6390 0.7322
-1.500 0.3823 0.00904 0.00372 -0.1104 0.6356 0.7461
-1.250 0.4094 0.00918 0.00386 -0.1101 0.6323 0.7560
-1.000 0.4371 0.00927 0.00392 -0.1100 0.6287 0.7640
-0.750 0.4642 0.00940 0.00403 -0.1097 0.6251 0.7716
-0.500 0.4916 0.00954 0.00412 -0.1095 0.6218 0.7804
-0.250 0.5180 0.00983 0.00437 -0.1089 0.6185 0.7914
0.000 0.5428 0.00994 0.00452 -0.1080 0.6152 0.7993
0.250 0.5695 0.01003 0.00460 -0.1077 0.6115 0.8047
0.500 0.5983 0.01007 0.00459 -0.1081 0.6080 0.8083
0.750 0.6277 0.01012 0.00457 -0.1087 0.6047 0.8110
1.000 0.6556 0.01018 0.00459 -0.1088 0.6013 0.8130
1.250 0.6827 0.01018 0.00462 -0.1088 0.5977 0.8150
1.500 0.7104 0.01021 0.00463 -0.1089 0.5940 0.8172
1.750 0.7385 0.01024 0.00464 -0.1091 0.5904 0.8195
2.000 0.7677 0.01034 0.00467 -0.1096 0.5870 0.8217
2.250 0.7958 0.01037 0.00472 -0.1099 0.5833 0.8241
2.500 0.8241 0.01041 0.00475 -0.1103 0.5792 0.8266
2.750 0.8513 0.01041 0.00476 -0.1103 0.5753 0.8285
3.000 0.8790 0.01048 0.00479 -0.1105 0.5715 0.8303
3.250 0.9058 0.01053 0.00487 -0.1105 0.5674 0.8323
3.500 0.9325 0.01056 0.00493 -0.1104 0.5629 0.8345
3.750 0.9597 0.01060 0.00497 -0.1105 0.5585 0.8369
4.000 0.9881 0.01073 0.00503 -0.1109 0.5543 0.8396
4.250 1.0146 0.01075 0.00512 -0.1109 0.5495 0.8421
4.500 1.0409 0.01078 0.00518 -0.1108 0.5447 0.8442
4.750 1.0672 0.01085 0.00523 -0.1107 0.5401 0.8462
5.000 1.0926 0.01092 0.00535 -0.1105 0.5352 0.8485
5.250 1.1180 0.01096 0.00545 -0.1103 0.5298 0.8510
5.500 1.1441 0.01106 0.00553 -0.1102 0.5247 0.8538
5.750 1.1697 0.01116 0.00566 -0.1100 0.5192 0.8569
6.000 1.1948 0.01122 0.00577 -0.1098 0.5131 0.8595
6.250 1.2191 0.01135 0.00588 -0.1093 0.5075 0.8620
6.500 1.2424 0.01140 0.00603 -0.1087 0.5010 0.8648
6.750 1.2658 0.01152 0.00617 -0.1081 0.4944 0.8679
7.000 1.2894 0.01165 0.00634 -0.1076 0.4875 0.8711
7.250 1.3126 0.01179 0.00650 -0.1070 0.4800 0.8744
7.500 1.3333 0.01191 0.00668 -0.1059 0.4723 0.8777
7.750 1.3533 0.01207 0.00687 -0.1047 0.4639 0.8816
8.000 1.3734 0.01224 0.00709 -0.1035 0.4552 0.8858
8.250 1.3908 0.01247 0.00730 -0.1019 0.4460 0.8901
8.500 1.4064 0.01262 0.00753 -0.0998 0.4361 0.8945
8.750 1.4210 0.01288 0.00782 -0.0976 0.4259 0.8999
9.000 1.4340 0.01322 0.00816 -0.0953 0.4144 0.9058
9.250 1.4457 0.01353 0.00852 -0.0926 0.4020 0.9127
9.500 1.4562 0.01393 0.00894 -0.0899 0.3881 0.9213
9.750 1.4636 0.01437 0.00940 -0.0867 0.3742 0.9327
10.000 1.4679 0.01487 0.00993 -0.0830 0.3593 0.9543
10.250 1.4787 0.01562 0.01066 -0.0813 0.3417 1.0000
10.750 1.4926 0.01772 0.01264 -0.0769 0.3060 1.0000
11.000 1.4969 0.01897 0.01384 -0.0746 0.2884 1.0000
11.250 1.4999 0.02036 0.01518 -0.0723 0.2719 1.0000
11.500 1.5020 0.02190 0.01668 -0.0701 0.2562 1.0000
11.750 1.5030 0.02359 0.01834 -0.0680 0.2407 1.0000
12.000 1.5038 0.02541 0.02012 -0.0661 0.2263 1.0000
12.250 1.5041 0.02734 0.02202 -0.0643 0.2124 1.0000
12.500 1.5037 0.02941 0.02406 -0.0626 0.1989 1.0000
12.750 1.5031 0.03159 0.02622 -0.0611 0.1863 1.0000
13.000 1.5017 0.03390 0.02850 -0.0597 0.1741 1.0000
13.250 1.5019 0.03617 0.03077 -0.0585 0.1625 1.0000
13.500 1.5032 0.03842 0.03302 -0.0575 0.1520 1.0000
13.750 1.5023 0.04092 0.03551 -0.0566 0.1413 1.0000
14.000 1.5010 0.04354 0.03812 -0.0557 0.1317 1.0000
14.250 1.5001 0.04621 0.04078 -0.0550 0.1218 1.0000
14.500 1.5001 0.04884 0.04342 -0.0544 0.1123 1.0000
14.750 1.4989 0.05168 0.04627 -0.0539 0.1037 1.0000
15.250 1.4972 0.05751 0.05211 -0.0533 0.0875 1.0000
15.500 1.4960 0.06056 0.05518 -0.0532 0.0806 1.0000
15.750 1.4935 0.06385 0.05847 -0.0532 0.0738 1.0000
16.000 1.4937 0.06688 0.06154 -0.0533 0.0673 1.0000
16.250 1.4908 0.07036 0.06502 -0.0535 0.0612 1.0000
16.500 1.4901 0.07364 0.06836 -0.0538 0.0567 1.0000
16.750 1.4879 0.07717 0.07190 -0.0542 0.0515 1.0000
17.000 1.4853 0.08083 0.07559 -0.0548 0.0470 1.0000
17.250 1.4833 0.08448 0.07930 -0.0554 0.0435 1.0000
17.500 1.4808 0.08823 0.08309 -0.0562 0.0397 1.0000
17.750 1.4779 0.09212 0.08705 -0.0570 0.0369 1.0000
18.000 1.4759 0.09594 0.09092 -0.0580 0.0337 1.0000
18.250 1.4706 0.10030 0.09532 -0.0592 0.0314 1.0000
18.500 1.4706 0.10389 0.09900 -0.0603 0.0289 1.0000
18.750 1.4653 0.10835 0.10350 -0.0617 0.0269 1.0000
19.000 1.4628 0.11241 0.10765 -0.0631 0.0251 1.0000
19.250 1.4606 0.11645 0.11176 -0.0646 0.0233 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 432 AIRFOIL (e432-il)