EPPLER 432 AIRFOIL (e432-il) Xfoil prediction polar at RE=231,000 Ncrit=9
| Details | Polar file |
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Airfoil: EPPLER 432 AIRFOIL (e432-il) Reynolds number: 231,000 Max Cl/Cd: 81.82 at α=9.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e432-il-231000.txt Download as CSV file: xf-e432-il-231000.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 = 0.231 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.1685 0.09961 0.09656 -0.0785 0.9627 0.0397
-10.000 -0.1564 0.09438 0.09133 -0.0847 0.9517 0.0420
-9.500 -0.1296 0.08124 0.07816 -0.1014 0.9326 0.0447
-9.250 -0.1022 0.07731 0.07418 -0.1068 0.9215 0.0461
-9.000 -0.0813 0.07135 0.06814 -0.1159 0.9066 0.0482
-8.750 -0.0903 0.06154 0.05812 -0.1309 0.8819 0.0505
-8.250 -0.1579 0.05622 0.05200 -0.1298 0.8379 0.0516
-8.000 -0.1455 0.05098 0.04683 -0.1304 0.8289 0.0526
-7.750 -0.1326 0.04873 0.04455 -0.1298 0.8192 0.0536
-7.500 -0.1252 0.04668 0.04241 -0.1289 0.8098 0.0549
-6.500 -0.1122 0.02893 0.02257 -0.1205 0.7810 0.0319
-6.250 -0.0959 0.02526 0.01798 -0.1178 0.7744 0.0257
-6.000 -0.0722 0.02326 0.01578 -0.1173 0.7691 0.0253
-5.750 -0.0482 0.02171 0.01399 -0.1166 0.7638 0.0251
-5.500 -0.0239 0.02045 0.01254 -0.1159 0.7579 0.0253
-5.250 0.0017 0.01928 0.01118 -0.1153 0.7529 0.0259
-5.000 0.0261 0.01832 0.01023 -0.1148 0.7482 0.0272
-4.750 0.0494 0.01763 0.00952 -0.1140 0.7428 0.0289
-4.500 0.0731 0.01693 0.00879 -0.1134 0.7380 0.0322
-4.250 0.0984 0.01641 0.00821 -0.1131 0.7339 0.0374
-4.000 0.1196 0.01575 0.00763 -0.1121 0.7287 0.0461
-3.750 0.1427 0.01511 0.00709 -0.1116 0.7239 0.0683
-3.500 0.1662 0.01432 0.00657 -0.1113 0.7199 0.1304
-3.250 0.1873 0.01316 0.00621 -0.1113 0.7160 0.3190
-3.000 0.2017 0.01216 0.00658 -0.1091 0.7110 0.6465
-2.750 0.2261 0.01254 0.00696 -0.1077 0.7066 0.7133
-2.500 0.2522 0.01296 0.00726 -0.1066 0.7029 0.7432
-2.250 0.2764 0.01341 0.00763 -0.1051 0.6993 0.7629
-2.000 0.2977 0.01380 0.00800 -0.1031 0.6945 0.7792
-1.750 0.3204 0.01413 0.00827 -0.1014 0.6902 0.7933
-1.500 0.3424 0.01444 0.00850 -0.0993 0.6867 0.8034
-1.250 0.3665 0.01471 0.00867 -0.0980 0.6834 0.8145
-1.000 0.3866 0.01494 0.00890 -0.0961 0.6787 0.8272
-0.750 0.4029 0.01521 0.00916 -0.0927 0.6745 0.8401
-0.500 0.4212 0.01540 0.00930 -0.0898 0.6710 0.8524
-0.250 0.4463 0.01550 0.00929 -0.0890 0.6680 0.8617
0.000 0.4663 0.01554 0.00934 -0.0874 0.6635 0.8674
0.250 0.4918 0.01556 0.00931 -0.0873 0.6591 0.8729
0.500 0.5182 0.01552 0.00920 -0.0872 0.6554 0.8765
0.750 0.5462 0.01552 0.00910 -0.0873 0.6523 0.8795
1.000 0.5700 0.01556 0.00914 -0.0869 0.6481 0.8832
1.250 0.5953 0.01560 0.00917 -0.0868 0.6436 0.8870
1.500 0.6240 0.01561 0.00913 -0.0875 0.6398 0.8902
1.750 0.6522 0.01559 0.00903 -0.0877 0.6365 0.8925
2.000 0.6769 0.01565 0.00909 -0.0873 0.6327 0.8956
2.250 0.7003 0.01571 0.00918 -0.0869 0.6280 0.8988
2.500 0.7283 0.01574 0.00918 -0.0873 0.6239 0.9017
2.750 0.7600 0.01578 0.00915 -0.0885 0.6205 0.9047
3.000 0.7847 0.01583 0.00921 -0.0881 0.6167 0.9074
3.250 0.8063 0.01589 0.00934 -0.0873 0.6117 0.9105
3.500 0.8333 0.01591 0.00934 -0.0875 0.6074 0.9135
3.750 0.8644 0.01592 0.00929 -0.0884 0.6039 0.9164
4.000 0.8895 0.01605 0.00947 -0.0883 0.5993 0.9200
4.250 0.9105 0.01606 0.00955 -0.0873 0.5941 0.9235
4.500 0.9381 0.01604 0.00951 -0.0875 0.5899 0.9267
4.750 0.9706 0.01607 0.00947 -0.0886 0.5863 0.9296
5.000 0.9896 0.01620 0.00973 -0.0875 0.5802 0.9340
5.250 1.0130 0.01615 0.00972 -0.0868 0.5753 0.9382
5.500 1.0433 0.01610 0.00962 -0.0875 0.5713 0.9422
5.750 1.0639 0.01622 0.00985 -0.0865 0.5655 0.9474
6.000 1.0871 0.01619 0.00987 -0.0859 0.5599 0.9522
6.250 1.1183 0.01613 0.00977 -0.0868 0.5555 0.9566
6.500 1.1391 0.01623 0.01000 -0.0859 0.5491 0.9630
6.750 1.1667 0.01620 0.01002 -0.0862 0.5431 0.9696
7.000 1.2011 0.01620 0.01001 -0.0879 0.5379 0.9753
7.250 1.2283 0.01632 0.01028 -0.0885 0.5301 0.9849
7.750 1.2843 0.01648 0.01054 -0.0898 0.5161 1.0000
8.000 1.3133 0.01650 0.01057 -0.0905 0.5091 1.0000
8.250 1.3350 0.01670 0.01088 -0.0901 0.5005 1.0000
8.500 1.3637 0.01677 0.01092 -0.0908 0.4927 1.0000
8.750 1.3823 0.01701 0.01130 -0.0899 0.4829 1.0000
9.000 1.4059 0.01719 0.01149 -0.0897 0.4738 1.0000
9.250 1.4244 0.01741 0.01177 -0.0886 0.4634 1.0000
9.500 1.4383 0.01773 0.01217 -0.0867 0.4524 1.0000
9.750 1.4535 0.01807 0.01253 -0.0851 0.4413 1.0000
10.000 1.4672 0.01848 0.01293 -0.0833 0.4294 1.0000
10.250 1.4766 0.01902 0.01351 -0.0809 0.4163 1.0000
10.500 1.4839 0.01969 0.01421 -0.0783 0.4023 1.0000
10.750 1.4896 0.02049 0.01503 -0.0757 0.3878 1.0000
11.000 1.4934 0.02145 0.01599 -0.0729 0.3726 1.0000
11.250 1.4956 0.02260 0.01712 -0.0702 0.3570 1.0000
11.500 1.4961 0.02395 0.01845 -0.0676 0.3409 1.0000
11.750 1.4950 0.02553 0.02000 -0.0650 0.3247 1.0000
12.000 1.4928 0.02730 0.02174 -0.0626 0.3086 1.0000
12.250 1.4899 0.02927 0.02368 -0.0604 0.2927 1.0000
12.500 1.4867 0.03139 0.02577 -0.0585 0.2773 1.0000
12.750 1.4832 0.03366 0.02801 -0.0567 0.2626 1.0000
13.000 1.4793 0.03608 0.03041 -0.0551 0.2481 1.0000
13.250 1.4759 0.03858 0.03291 -0.0538 0.2341 1.0000
13.500 1.4724 0.04120 0.03553 -0.0526 0.2203 1.0000
13.750 1.4692 0.04391 0.03825 -0.0516 0.2071 1.0000
14.000 1.4657 0.04673 0.04107 -0.0508 0.1943 1.0000
14.250 1.4625 0.04964 0.04398 -0.0501 0.1825 1.0000
14.500 1.4585 0.05272 0.04706 -0.0496 0.1710 1.0000
14.750 1.4535 0.05599 0.05030 -0.0492 0.1601 1.0000
15.000 1.4506 0.05916 0.05350 -0.0490 0.1491 1.0000
15.250 1.4482 0.06237 0.05674 -0.0489 0.1387 1.0000
15.500 1.4445 0.06581 0.06019 -0.0489 0.1292 1.0000
15.750 1.4385 0.06958 0.06394 -0.0490 0.1206 1.0000
16.000 1.4367 0.07297 0.06739 -0.0493 0.1114 1.0000
16.250 1.4331 0.07664 0.07109 -0.0497 0.1034 1.0000
16.500 1.4272 0.08067 0.07510 -0.0502 0.0962 1.0000
16.750 1.4252 0.08429 0.07881 -0.0508 0.0887 1.0000
17.000 1.4190 0.08849 0.08299 -0.0516 0.0827 1.0000
17.250 1.4170 0.09225 0.08684 -0.0524 0.0764 1.0000
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Polar data table (+)
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