EPPLER 417 AIRFOIL (e417-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 417 AIRFOIL (e417-il) Reynolds number: 200,000 Max Cl/Cd: 66.14 at α=4.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e417-il-200000-n5.txt Download as CSV file: xf-e417-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 417 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
------ -------- --------- --------- -------- -------- --------
-11.750 -0.4949 0.10518 0.10156 -0.0579 0.9901 0.0159
-11.500 -0.5809 0.07595 0.07199 -0.0767 0.9875 0.0141
-11.250 -0.6044 0.06864 0.06438 -0.0822 0.9856 0.0138
-10.750 -0.6348 0.06052 0.05587 -0.0847 0.9790 0.0136
-10.500 -0.6468 0.05659 0.05173 -0.0860 0.9762 0.0136
-10.250 -0.6548 0.05338 0.04834 -0.0866 0.9736 0.0135
-10.000 -0.6701 0.05113 0.04597 -0.0840 0.9681 0.0135
-9.750 -0.6763 0.04824 0.04285 -0.0838 0.9645 0.0134
-9.500 -0.6686 0.04497 0.03927 -0.0853 0.9623 0.0134
-9.250 -0.6755 0.04268 0.03673 -0.0821 0.9566 0.0134
-9.000 -0.6659 0.04009 0.03384 -0.0817 0.9538 0.0134
-8.750 -0.6493 0.03754 0.03093 -0.0822 0.9517 0.0134
-8.500 -0.6280 0.03535 0.02843 -0.0829 0.9503 0.0134
-8.250 -0.6138 0.03374 0.02656 -0.0817 0.9477 0.0135
-8.000 -0.6010 0.03220 0.02482 -0.0799 0.9444 0.0137
-7.750 -0.5811 0.03057 0.02304 -0.0793 0.9422 0.0138
-7.500 -0.5591 0.02916 0.02154 -0.0790 0.9405 0.0141
-7.250 -0.5355 0.02801 0.02030 -0.0790 0.9391 0.0144
-7.000 -0.5100 0.02698 0.01920 -0.0794 0.9378 0.0148
-6.750 -0.4821 0.02613 0.01830 -0.0803 0.9367 0.0159
-6.500 -0.4670 0.02541 0.01752 -0.0787 0.9332 0.0165
-6.250 -0.4477 0.02464 0.01669 -0.0779 0.9300 0.0171
-6.000 -0.4226 0.02379 0.01576 -0.0783 0.9277 0.0174
-5.750 -0.3945 0.02297 0.01486 -0.0794 0.9259 0.0179
-5.500 -0.3640 0.02211 0.01394 -0.0811 0.9246 0.0185
-5.250 -0.3315 0.02135 0.01310 -0.0832 0.9235 0.0195
-5.000 -0.2977 0.02079 0.01244 -0.0852 0.9226 0.0210
-4.750 -0.2793 0.02044 0.01202 -0.0841 0.9181 0.0226
-4.500 -0.2510 0.01995 0.01152 -0.0850 0.9154 0.0281
-4.250 -0.2184 0.01928 0.01096 -0.0870 0.9135 0.0534
-4.000 -0.1764 0.01740 0.01006 -0.0929 0.9131 0.2622
-3.750 -0.1377 0.01637 0.01015 -0.0968 0.9123 0.5133
-3.500 -0.1059 0.01669 0.01052 -0.0974 0.9106 0.5672
-3.250 -0.0741 0.01720 0.01100 -0.0977 0.9090 0.5966
-3.000 -0.0548 0.01759 0.01137 -0.0958 0.9042 0.6096
-2.750 -0.0293 0.01771 0.01144 -0.0953 0.9005 0.6140
-2.500 0.0018 0.01771 0.01136 -0.0961 0.8982 0.6172
-2.250 0.0354 0.01764 0.01120 -0.0975 0.8964 0.6209
-2.000 0.0716 0.01750 0.01093 -0.0995 0.8951 0.6252
-1.750 0.0935 0.01756 0.01096 -0.0986 0.8900 0.6275
-1.500 0.1208 0.01754 0.01092 -0.0986 0.8860 0.6296
-1.250 0.1531 0.01746 0.01080 -0.0996 0.8835 0.6322
-1.000 0.1879 0.01733 0.01064 -0.1012 0.8818 0.6351
-0.750 0.2245 0.01715 0.01041 -0.1032 0.8804 0.6382
-0.500 0.2433 0.01723 0.01044 -0.1018 0.8725 0.6416
-0.250 0.2754 0.01709 0.01032 -0.1027 0.8695 0.6434
0.000 0.3106 0.01688 0.01012 -0.1041 0.8675 0.6454
0.500 0.3646 0.01674 0.01000 -0.1040 0.8565 0.6504
0.750 0.4008 0.01643 0.00970 -0.1057 0.8539 0.6534
1.000 0.4389 0.01605 0.00930 -0.1077 0.8515 0.6563
1.250 0.4580 0.01609 0.00940 -0.1061 0.8419 0.6582
1.500 0.4919 0.01578 0.00915 -0.1071 0.8390 0.6602
1.750 0.5142 0.01580 0.00923 -0.1061 0.8307 0.6625
2.000 0.5461 0.01554 0.00902 -0.1069 0.8258 0.6648
2.250 0.5727 0.01539 0.00891 -0.1066 0.8168 0.6674
2.500 0.6081 0.01497 0.00854 -0.1079 0.8095 0.6703
2.750 0.6345 0.01485 0.00846 -0.1076 0.7975 0.6729
3.000 0.6641 0.01464 0.00834 -0.1078 0.7863 0.6748
3.250 0.6980 0.01432 0.00809 -0.1087 0.7736 0.6768
3.500 0.7359 0.01393 0.00775 -0.1104 0.7569 0.6789
3.750 0.7823 0.01344 0.00727 -0.1138 0.7339 0.6814
4.000 0.8279 0.01315 0.00692 -0.1171 0.6992 0.6842
4.250 0.8681 0.01317 0.00675 -0.1195 0.6527 0.6871
4.500 0.8936 0.01351 0.00689 -0.1189 0.6022 0.6890
4.750 0.9107 0.01400 0.00719 -0.1168 0.5533 0.6909
5.000 0.9228 0.01462 0.00758 -0.1137 0.4989 0.6931
5.250 0.9310 0.01541 0.00805 -0.1101 0.4365 0.6956
5.500 0.9379 0.01634 0.00861 -0.1065 0.3680 0.6985
5.750 0.9485 0.01725 0.00919 -0.1038 0.3052 0.7015
6.000 0.9619 0.01807 0.00978 -0.1016 0.2532 0.7038
6.250 0.9759 0.01889 0.01039 -0.0995 0.2055 0.7058
6.500 0.9897 0.01983 0.01107 -0.0975 0.1551 0.7080
6.750 1.0047 0.02075 0.01176 -0.0958 0.1117 0.7105
7.000 1.0198 0.02173 0.01252 -0.0941 0.0737 0.7133
7.250 1.0356 0.02269 0.01335 -0.0926 0.0445 0.7164
7.500 1.0509 0.02368 0.01425 -0.0909 0.0266 0.7190
7.750 1.0670 0.02454 0.01515 -0.0892 0.0208 0.7212
8.000 1.0837 0.02537 0.01606 -0.0876 0.0183 0.7236
8.250 1.1003 0.02621 0.01699 -0.0861 0.0164 0.7263
8.500 1.1157 0.02719 0.01806 -0.0846 0.0152 0.7292
8.750 1.1315 0.02812 0.01912 -0.0831 0.0144 0.7324
9.000 1.1456 0.02912 0.02027 -0.0813 0.0138 0.7350
9.250 1.1585 0.03020 0.02149 -0.0794 0.0134 0.7377
9.500 1.1706 0.03138 0.02279 -0.0775 0.0130 0.7406
9.750 1.1820 0.03265 0.02418 -0.0756 0.0127 0.7436
10.000 1.1923 0.03406 0.02571 -0.0737 0.0123 0.7466
10.250 1.2016 0.03558 0.02735 -0.0718 0.0121 0.7494
10.500 1.2096 0.03725 0.02914 -0.0697 0.0118 0.7520
10.750 1.2175 0.03913 0.03113 -0.0678 0.0116 0.7550
11.000 1.2276 0.04091 0.03304 -0.0662 0.0115 0.7585
11.250 1.2387 0.04264 0.03492 -0.0647 0.0113 0.7620
11.500 1.2489 0.04442 0.03690 -0.0632 0.0111 0.7650
11.750 1.2582 0.04631 0.03899 -0.0617 0.0108 0.7680
12.000 1.2662 0.04833 0.04124 -0.0603 0.0106 0.7713
12.250 1.2728 0.05051 0.04363 -0.0589 0.0102 0.7749
12.500 1.2781 0.05290 0.04622 -0.0576 0.0100 0.7786
12.750 1.2811 0.05550 0.04905 -0.0561 0.0099 0.7818
13.000 1.2821 0.05835 0.05214 -0.0548 0.0098 0.7855
13.250 1.2810 0.06149 0.05552 -0.0537 0.0097 0.7894
13.750 1.2716 0.06868 0.06321 -0.0521 0.0096 0.7966
14.000 1.2636 0.07282 0.06761 -0.0518 0.0095 0.8002
14.250 1.2537 0.07742 0.07247 -0.0521 0.0095 0.8040
14.500 1.2417 0.08257 0.07787 -0.0530 0.0095 0.8079
14.750 1.2275 0.08823 0.08380 -0.0545 0.0095 0.8113
15.000 1.2117 0.09457 0.09041 -0.0568 0.0095 0.8153
15.250 1.1939 0.10173 0.09783 -0.0601 0.0095 0.8194
15.750 1.1531 0.11875 0.11537 -0.0700 0.0096 0.8269
16.000 1.1287 0.12935 0.12622 -0.0770 0.0097 0.8302
16.250 1.1002 0.14219 0.13931 -0.0858 0.0098 0.8324
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 417 AIRFOIL (e417-il)