EPPLER 432 AIRFOIL (e432-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 432 AIRFOIL (e432-il) Reynolds number: 1,000,000 Max Cl/Cd: 131.63 at α=5.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e432-il-1000000-n5.txt Download as CSV file: xf-e432-il-1000000-n5.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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.500 -0.4164 0.04062 0.03761 -0.1371 0.8195 0.0041
-12.250 -0.4448 0.03600 0.03268 -0.1374 0.7923 0.0041
-12.000 -0.4515 0.03379 0.03028 -0.1366 0.7739 0.0042
-11.750 -0.4724 0.03059 0.02681 -0.1347 0.7594 0.0041
-11.500 -0.4792 0.02867 0.02469 -0.1326 0.7472 0.0042
-11.250 -0.4856 0.02686 0.02267 -0.1300 0.7359 0.0042
-11.000 -0.4894 0.02538 0.02102 -0.1269 0.7266 0.0042
-10.500 -0.4827 0.02310 0.01839 -0.1217 0.7096 0.0042
-10.250 -0.4735 0.02187 0.01696 -0.1198 0.7020 0.0043
-10.000 -0.4594 0.02088 0.01582 -0.1184 0.6961 0.0043
-9.750 -0.4447 0.01982 0.01459 -0.1170 0.6899 0.0043
-9.500 -0.4276 0.01897 0.01358 -0.1158 0.6836 0.0043
-9.250 -0.4089 0.01813 0.01260 -0.1149 0.6783 0.0044
-9.000 -0.3886 0.01747 0.01182 -0.1141 0.6728 0.0044
-8.500 -0.3470 0.01603 0.01014 -0.1125 0.6638 0.0044
-8.250 -0.3247 0.01544 0.00945 -0.1119 0.6590 0.0045
-8.000 -0.3011 0.01503 0.00895 -0.1115 0.6544 0.0045
-7.750 -0.2774 0.01460 0.00843 -0.1110 0.6501 0.0046
-7.500 -0.2532 0.01415 0.00792 -0.1107 0.6464 0.0046
-7.250 -0.2309 0.01342 0.00710 -0.1100 0.6424 0.0048
-7.000 -0.2069 0.01295 0.00655 -0.1096 0.6384 0.0049
-6.750 -0.1824 0.01256 0.00609 -0.1092 0.6344 0.0050
-6.500 -0.1569 0.01218 0.00567 -0.1090 0.6312 0.0051
-6.250 -0.1312 0.01185 0.00529 -0.1088 0.6279 0.0052
-6.000 -0.1053 0.01155 0.00495 -0.1087 0.6243 0.0053
-5.750 -0.0794 0.01128 0.00462 -0.1085 0.6207 0.0054
-5.500 -0.0531 0.01102 0.00432 -0.1084 0.6174 0.0055
-5.250 -0.0262 0.01077 0.00404 -0.1084 0.6144 0.0058
-5.000 0.0008 0.01054 0.00378 -0.1083 0.6113 0.0060
-4.750 0.0278 0.01034 0.00355 -0.1083 0.6078 0.0063
-4.500 0.0548 0.01016 0.00333 -0.1083 0.6045 0.0067
-4.250 0.0820 0.00998 0.00312 -0.1083 0.6015 0.0075
-4.000 0.1097 0.00980 0.00293 -0.1084 0.5988 0.0092
-3.750 0.1375 0.00964 0.00277 -0.1086 0.5958 0.0112
-3.500 0.1651 0.00948 0.00262 -0.1087 0.5924 0.0159
-3.250 0.1926 0.00934 0.00249 -0.1088 0.5891 0.0239
-3.000 0.2202 0.00921 0.00238 -0.1089 0.5861 0.0343
-2.750 0.2484 0.00903 0.00227 -0.1092 0.5835 0.0515
-2.500 0.2763 0.00881 0.00217 -0.1095 0.5804 0.0840
-2.250 0.3042 0.00856 0.00208 -0.1098 0.5771 0.1303
-2.000 0.3320 0.00828 0.00199 -0.1102 0.5740 0.1966
-1.750 0.3598 0.00802 0.00191 -0.1106 0.5707 0.2649
-1.500 0.3885 0.00772 0.00184 -0.1111 0.5680 0.3450
-1.250 0.4175 0.00719 0.00173 -0.1120 0.5648 0.4876
-1.000 0.4466 0.00677 0.00172 -0.1127 0.5616 0.6291
-0.750 0.4752 0.00676 0.00175 -0.1130 0.5582 0.6638
-0.500 0.5035 0.00680 0.00178 -0.1132 0.5548 0.6812
-0.250 0.5325 0.00683 0.00181 -0.1135 0.5519 0.6949
0.000 0.5613 0.00687 0.00186 -0.1137 0.5487 0.7087
0.250 0.5897 0.00693 0.00193 -0.1139 0.5450 0.7242
0.500 0.6177 0.00700 0.00200 -0.1140 0.5413 0.7349
0.750 0.6460 0.00708 0.00204 -0.1142 0.5380 0.7402
1.000 0.6749 0.00712 0.00207 -0.1145 0.5347 0.7423
1.250 0.7035 0.00717 0.00209 -0.1148 0.5308 0.7441
1.500 0.7315 0.00723 0.00213 -0.1150 0.5268 0.7459
1.750 0.7592 0.00731 0.00218 -0.1151 0.5227 0.7478
2.000 0.7876 0.00735 0.00223 -0.1153 0.5187 0.7497
2.250 0.8156 0.00742 0.00229 -0.1155 0.5139 0.7515
2.500 0.8430 0.00751 0.00235 -0.1156 0.5089 0.7534
2.750 0.8709 0.00758 0.00241 -0.1158 0.5043 0.7552
3.000 0.8986 0.00766 0.00248 -0.1159 0.4987 0.7571
3.250 0.9256 0.00777 0.00256 -0.1160 0.4931 0.7591
3.500 0.9532 0.00786 0.00264 -0.1161 0.4878 0.7609
3.750 0.9802 0.00794 0.00272 -0.1161 0.4817 0.7627
4.000 1.0065 0.00806 0.00283 -0.1160 0.4757 0.7645
4.250 1.0336 0.00815 0.00293 -0.1161 0.4696 0.7663
4.500 1.0595 0.00829 0.00305 -0.1159 0.4627 0.7682
4.750 1.0860 0.00840 0.00316 -0.1159 0.4561 0.7702
5.000 1.1116 0.00855 0.00330 -0.1157 0.4482 0.7722
5.250 1.1373 0.00869 0.00343 -0.1155 0.4408 0.7744
5.500 1.1621 0.00887 0.00358 -0.1152 0.4320 0.7765
5.750 1.1873 0.00902 0.00374 -0.1150 0.4236 0.7784
6.000 1.2107 0.00923 0.00392 -0.1144 0.4134 0.7804
6.250 1.2344 0.00942 0.00411 -0.1139 0.4027 0.7825
6.500 1.2574 0.00964 0.00431 -0.1132 0.3918 0.7846
6.750 1.2788 0.00992 0.00455 -0.1123 0.3782 0.7869
7.000 1.2993 0.01021 0.00480 -0.1112 0.3643 0.7893
7.250 1.3175 0.01054 0.00508 -0.1097 0.3486 0.7917
7.500 1.3338 0.01088 0.00537 -0.1078 0.3334 0.7941
7.750 1.3483 0.01128 0.00572 -0.1056 0.3166 0.7969
8.000 1.3619 0.01173 0.00612 -0.1033 0.2994 0.7998
8.250 1.3752 0.01221 0.00654 -0.1010 0.2835 0.8028
8.500 1.3870 0.01274 0.00702 -0.0986 0.2669 0.8059
8.750 1.3976 0.01334 0.00755 -0.0960 0.2501 0.8089
9.000 1.4081 0.01394 0.00811 -0.0934 0.2352 0.8118
9.250 1.4180 0.01458 0.00872 -0.0909 0.2210 0.8151
9.500 1.4269 0.01529 0.00940 -0.0883 0.2072 0.8189
9.750 1.4351 0.01608 0.01015 -0.0858 0.1938 0.8230
10.250 1.4484 0.01793 0.01194 -0.0807 0.1679 0.8314
10.500 1.4560 0.01889 0.01291 -0.0785 0.1587 0.8360
10.750 1.4603 0.02010 0.01410 -0.0761 0.1464 0.8409
11.000 1.4659 0.02131 0.01530 -0.0739 0.1358 0.8458
11.250 1.4684 0.02278 0.01675 -0.0717 0.1234 0.8518
11.500 1.4731 0.02419 0.01817 -0.0698 0.1140 0.8581
11.750 1.4776 0.02567 0.01967 -0.0680 0.1063 0.8657
12.250 1.4826 0.02907 0.02310 -0.0645 0.0878 0.8854
12.500 1.4812 0.03110 0.02514 -0.0626 0.0771 0.9002
12.750 1.4822 0.03282 0.02695 -0.0607 0.0710 0.9316
13.250 1.4910 0.03672 0.03091 -0.0591 0.0587 1.0000
13.500 1.4858 0.03959 0.03371 -0.0579 0.0469 1.0000
13.750 1.4911 0.04157 0.03573 -0.0573 0.0441 1.0000
14.000 1.4930 0.04394 0.03809 -0.0566 0.0389 1.0000
14.250 1.4973 0.04612 0.04030 -0.0561 0.0353 1.0000
14.500 1.4996 0.04856 0.04275 -0.0556 0.0318 1.0000
14.750 1.5031 0.05093 0.04516 -0.0552 0.0286 1.0000
15.000 1.5054 0.05347 0.04772 -0.0549 0.0253 1.0000
15.250 1.5088 0.05596 0.05026 -0.0548 0.0232 1.0000
15.500 1.5114 0.05859 0.05292 -0.0547 0.0210 1.0000
15.750 1.5148 0.06115 0.05552 -0.0547 0.0190 1.0000
16.000 1.5165 0.06399 0.05840 -0.0547 0.0173 1.0000
16.250 1.5180 0.06690 0.06134 -0.0549 0.0150 1.0000
16.500 1.5211 0.06967 0.06417 -0.0551 0.0143 1.0000
16.750 1.5224 0.07270 0.06725 -0.0554 0.0126 1.0000
17.000 1.5238 0.07576 0.07037 -0.0558 0.0115 1.0000
17.250 1.5252 0.07889 0.07356 -0.0563 0.0105 1.0000
17.500 1.5253 0.08224 0.07696 -0.0569 0.0093 1.0000
17.750 1.5263 0.08552 0.08030 -0.0576 0.0086 1.0000
18.000 1.5254 0.08909 0.08393 -0.0584 0.0075 1.0000
18.250 1.5260 0.09247 0.08738 -0.0592 0.0070 1.0000
18.500 1.5256 0.09609 0.09107 -0.0602 0.0063 1.0000
18.750 1.5250 0.09975 0.09479 -0.0612 0.0058 1.0000
19.000 1.5239 0.10353 0.09865 -0.0624 0.0053 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 432 AIRFOIL (e432-il)