EPPLER 432 AIRFOIL (e432-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 432 AIRFOIL (e432-il) Reynolds number: 1,000,000 Max Cl/Cd: 141.36 at α=6.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e432-il-1000000.txt Download as CSV file: xf-e432-il-1000000.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 = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.000 -0.0466 0.07323 0.07072 -0.1079 0.7730 0.0116
-10.750 -0.0437 0.07060 0.06805 -0.1085 0.7622 0.0118
-10.500 -0.0414 0.06768 0.06508 -0.1093 0.7520 0.0120
-10.250 -0.0445 0.06331 0.06070 -0.1109 0.7442 0.0121
-10.000 -0.0494 0.05865 0.05601 -0.1126 0.7367 0.0122
-9.750 -0.0634 0.05193 0.04929 -0.1156 0.7310 0.0122
-9.500 -0.1499 0.03145 0.02864 -0.1285 0.7308 0.0119
-9.250 -0.1874 0.02648 0.02346 -0.1280 0.7237 0.0119
-9.000 -0.1798 0.02698 0.02396 -0.1270 0.7157 0.0121
-8.750 -0.2116 0.02368 0.02047 -0.1231 0.7090 0.0121
-8.500 -0.2972 0.02391 0.01941 -0.1197 0.7151 0.0084
-8.250 -0.2878 0.02152 0.01669 -0.1178 0.7083 0.0083
-8.000 -0.2716 0.02007 0.01498 -0.1165 0.7021 0.0083
-7.750 -0.2537 0.01848 0.01316 -0.1154 0.6968 0.0082
-7.500 -0.2333 0.01723 0.01171 -0.1145 0.6914 0.0082
-7.250 -0.2121 0.01611 0.01038 -0.1137 0.6861 0.0081
-7.000 -0.1891 0.01525 0.00939 -0.1131 0.6815 0.0081
-6.750 -0.1656 0.01447 0.00850 -0.1125 0.6767 0.0081
-6.500 -0.1415 0.01389 0.00782 -0.1121 0.6722 0.0083
-6.250 -0.1168 0.01350 0.00732 -0.1116 0.6676 0.0085
-6.000 -0.0917 0.01297 0.00674 -0.1113 0.6640 0.0086
-5.750 -0.0668 0.01246 0.00617 -0.1110 0.6599 0.0087
-5.500 -0.0420 0.01199 0.00564 -0.1106 0.6558 0.0088
-5.250 -0.0170 0.01161 0.00517 -0.1102 0.6517 0.0088
-5.000 0.0082 0.01109 0.00461 -0.1100 0.6486 0.0090
-4.750 0.0339 0.01062 0.00410 -0.1098 0.6449 0.0093
-4.500 0.0604 0.01031 0.00376 -0.1097 0.6413 0.0099
-4.250 0.0872 0.01009 0.00350 -0.1096 0.6376 0.0105
-4.000 0.1141 0.00983 0.00319 -0.1096 0.6341 0.0128
-3.750 0.1416 0.00957 0.00294 -0.1096 0.6311 0.0174
-3.500 0.1689 0.00926 0.00273 -0.1097 0.6278 0.0364
-3.250 0.1964 0.00905 0.00259 -0.1099 0.6243 0.0562
-3.000 0.2238 0.00886 0.00247 -0.1101 0.6209 0.0842
-2.750 0.2514 0.00860 0.00236 -0.1104 0.6176 0.1320
-2.250 0.3066 0.00740 0.00207 -0.1117 0.6115 0.4036
-2.000 0.3348 0.00660 0.00198 -0.1125 0.6083 0.6350
-1.750 0.3634 0.00661 0.00202 -0.1127 0.6049 0.6760
-1.500 0.3922 0.00667 0.00205 -0.1130 0.6015 0.6946
-1.250 0.4213 0.00670 0.00208 -0.1132 0.5987 0.7103
-1.000 0.4500 0.00674 0.00212 -0.1134 0.5956 0.7213
-0.750 0.4789 0.00680 0.00214 -0.1136 0.5923 0.7296
-0.500 0.5071 0.00688 0.00219 -0.1137 0.5889 0.7364
-0.250 0.5359 0.00699 0.00224 -0.1140 0.5855 0.7433
0.000 0.5643 0.00702 0.00231 -0.1141 0.5827 0.7504
0.250 0.5927 0.00713 0.00240 -0.1142 0.5794 0.7596
0.500 0.6203 0.00720 0.00249 -0.1141 0.5761 0.7672
0.750 0.6482 0.00731 0.00256 -0.1142 0.5726 0.7722
1.000 0.6768 0.00738 0.00260 -0.1145 0.5691 0.7747
1.250 0.7057 0.00740 0.00262 -0.1148 0.5659 0.7767
1.500 0.7344 0.00743 0.00263 -0.1151 0.5624 0.7785
1.750 0.7623 0.00746 0.00264 -0.1153 0.5586 0.7805
2.000 0.7899 0.00755 0.00269 -0.1154 0.5543 0.7823
2.250 0.8184 0.00757 0.00273 -0.1157 0.5509 0.7842
2.500 0.8465 0.00760 0.00278 -0.1159 0.5468 0.7862
2.750 0.8742 0.00767 0.00282 -0.1160 0.5425 0.7882
3.000 0.9017 0.00777 0.00289 -0.1162 0.5380 0.7902
3.250 0.9301 0.00780 0.00294 -0.1165 0.5339 0.7920
3.500 0.9579 0.00787 0.00300 -0.1167 0.5291 0.7937
3.750 0.9846 0.00795 0.00306 -0.1166 0.5243 0.7958
4.000 1.0121 0.00799 0.00313 -0.1168 0.5199 0.7980
4.250 1.0394 0.00805 0.00321 -0.1169 0.5148 0.7999
4.500 1.0658 0.00816 0.00330 -0.1168 0.5096 0.8020
4.750 1.0928 0.00824 0.00340 -0.1168 0.5046 0.8040
5.000 1.1198 0.00832 0.00350 -0.1169 0.4988 0.8062
5.250 1.1453 0.00846 0.00361 -0.1167 0.4929 0.8083
5.500 1.1725 0.00854 0.00371 -0.1168 0.4873 0.8102
5.750 1.1982 0.00864 0.00382 -0.1166 0.4808 0.8126
6.000 1.2232 0.00877 0.00396 -0.1163 0.4743 0.8150
6.250 1.2487 0.00888 0.00409 -0.1161 0.4672 0.8173
6.500 1.2728 0.00904 0.00425 -0.1157 0.4599 0.8198
6.750 1.2977 0.00918 0.00440 -0.1154 0.4514 0.8223
7.000 1.3212 0.00936 0.00458 -0.1148 0.4430 0.8247
7.250 1.3445 0.00956 0.00476 -0.1143 0.4332 0.8269
7.500 1.3672 0.00972 0.00495 -0.1136 0.4235 0.8298
7.750 1.3878 0.00996 0.00518 -0.1125 0.4123 0.8326
8.000 1.4063 0.01023 0.00543 -0.1110 0.4001 0.8358
8.250 1.4241 0.01049 0.00568 -0.1094 0.3869 0.8392
8.500 1.4396 0.01085 0.00599 -0.1074 0.3715 0.8425
8.750 1.4537 0.01123 0.00634 -0.1052 0.3555 0.8462
9.000 1.4661 0.01169 0.00676 -0.1027 0.3389 0.8501
9.250 1.4779 0.01220 0.00722 -0.1002 0.3209 0.8543
9.500 1.4880 0.01279 0.00775 -0.0975 0.3019 0.8586
9.750 1.4955 0.01345 0.00836 -0.0944 0.2833 0.8636
10.000 1.5024 0.01417 0.00904 -0.0914 0.2663 0.8695
10.250 1.5093 0.01494 0.00977 -0.0886 0.2508 0.8756
10.500 1.5141 0.01580 0.01061 -0.0855 0.2353 0.8829
10.750 1.5173 0.01681 0.01158 -0.0824 0.2193 0.8913
11.000 1.5205 0.01785 0.01262 -0.0794 0.2051 0.9023
11.250 1.5219 0.01899 0.01376 -0.0763 0.1912 0.9189
11.500 1.5237 0.02012 0.01496 -0.0736 0.1781 1.0000
11.750 1.5290 0.02150 0.01631 -0.0718 0.1662 1.0000
12.000 1.5312 0.02314 0.01790 -0.0698 0.1528 1.0000
12.250 1.5331 0.02488 0.01960 -0.0679 0.1404 1.0000
12.500 1.5325 0.02688 0.02154 -0.0660 0.1265 1.0000
12.750 1.5340 0.02880 0.02344 -0.0645 0.1163 1.0000
13.000 1.5352 0.03083 0.02544 -0.0630 0.1071 1.0000
13.250 1.5359 0.03295 0.02754 -0.0617 0.0974 1.0000
13.500 1.5335 0.03542 0.02996 -0.0603 0.0859 1.0000
13.750 1.5370 0.03745 0.03201 -0.0594 0.0795 1.0000
14.000 1.5372 0.03983 0.03437 -0.0584 0.0723 1.0000
14.250 1.5396 0.04208 0.03664 -0.0577 0.0662 1.0000
14.500 1.5413 0.04446 0.03903 -0.0570 0.0609 1.0000
14.750 1.5413 0.04707 0.04164 -0.0564 0.0549 1.0000
15.000 1.5433 0.04953 0.04411 -0.0559 0.0496 1.0000
15.250 1.5431 0.05229 0.04689 -0.0555 0.0449 1.0000
15.500 1.5455 0.05485 0.04948 -0.0553 0.0412 1.0000
15.750 1.5462 0.05764 0.05228 -0.0551 0.0375 1.0000
16.000 1.5481 0.06035 0.05503 -0.0550 0.0344 1.0000
16.250 1.5475 0.06341 0.05811 -0.0550 0.0309 1.0000
16.500 1.5506 0.06609 0.06085 -0.0551 0.0288 1.0000
16.750 1.5490 0.06941 0.06420 -0.0553 0.0259 1.0000
17.000 1.5507 0.07237 0.06721 -0.0556 0.0238 1.0000
17.250 1.5496 0.07572 0.07059 -0.0560 0.0218 1.0000
17.500 1.5506 0.07888 0.07382 -0.0565 0.0201 1.0000
17.750 1.5499 0.08231 0.07728 -0.0571 0.0184 1.0000
18.000 1.5498 0.08572 0.08076 -0.0578 0.0170 1.0000
18.250 1.5492 0.08922 0.08432 -0.0586 0.0157 1.0000
18.500 1.5475 0.09293 0.08808 -0.0594 0.0145 1.0000
18.750 1.5471 0.09652 0.09174 -0.0604 0.0136 1.0000
19.000 1.5457 0.10028 0.09558 -0.0615 0.0127 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 432 AIRFOIL (e432-il)