EPPLER 66 AIRFOIL (e66-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: EPPLER 66 AIRFOIL (e66-il) Reynolds number: 200,000 Max Cl/Cd: 91.72 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e66-il-200000.txt Download as CSV file: xf-e66-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 66 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.4222 0.10001 0.09688 -0.0331 1.0000 0.0496
-8.000 -0.4389 0.09806 0.09500 -0.0320 1.0000 0.0497
-7.750 -0.4568 0.09628 0.09329 -0.0302 1.0000 0.0498
-7.500 -0.4677 0.09334 0.09040 -0.0313 0.9995 0.0499
-7.250 -0.4598 0.08704 0.08414 -0.0326 0.9966 0.0511
-7.000 -0.4397 0.08460 0.08167 -0.0324 0.9942 0.0528
-6.750 -0.4220 0.08050 0.07757 -0.0373 0.9888 0.0545
-6.500 -0.3976 0.07432 0.07135 -0.0478 0.9837 0.0575
-6.250 -0.3718 0.05782 0.05441 -0.0759 0.9731 0.0634
-6.000 -0.3524 0.04209 0.03801 -0.0845 0.9682 0.0368
-5.750 -0.3225 0.03390 0.02904 -0.0906 0.9644 0.0336
-5.500 -0.2857 0.02958 0.02404 -0.0945 0.9620 0.0346
-5.250 -0.2467 0.02574 0.01947 -0.0976 0.9605 0.0343
-5.000 -0.2075 0.02354 0.01678 -0.1002 0.9590 0.0358
-4.750 -0.1807 0.02235 0.01529 -0.1000 0.9531 0.0382
-4.500 -0.1457 0.02081 0.01367 -0.1019 0.9502 0.0438
-4.250 -0.1068 0.02004 0.01269 -0.1040 0.9477 0.0499
-4.000 -0.0672 0.01892 0.01156 -0.1066 0.9460 0.0609
-3.750 -0.0420 0.01821 0.01086 -0.1064 0.9392 0.0742
-3.500 -0.0047 0.01733 0.01010 -0.1085 0.9360 0.1060
-3.250 0.0353 0.01674 0.00966 -0.1113 0.9336 0.1523
-3.000 0.0652 0.01639 0.00944 -0.1120 0.9281 0.1971
-2.750 0.0990 0.01600 0.00924 -0.1135 0.9234 0.2546
-2.500 0.1382 0.01558 0.00901 -0.1160 0.9207 0.3213
-2.250 0.1791 0.01516 0.00879 -0.1187 0.9187 0.3917
-2.000 0.2026 0.01498 0.00878 -0.1180 0.9105 0.4501
-1.750 0.2404 0.01460 0.00862 -0.1199 0.9074 0.5245
-1.500 0.2802 0.01419 0.00841 -0.1220 0.9053 0.5989
-1.250 0.3040 0.01405 0.00845 -0.1210 0.8976 0.6597
-1.000 0.3386 0.01370 0.00826 -0.1218 0.8940 0.7255
-0.750 0.3749 0.01327 0.00798 -0.1227 0.8917 0.7889
-0.500 0.3943 0.01310 0.00792 -0.1203 0.8841 0.8464
-0.250 0.4230 0.01268 0.00759 -0.1195 0.8798 0.9188
0.000 0.4737 0.01226 0.00711 -0.1238 0.8778 1.0000
0.250 0.5073 0.01215 0.00692 -0.1250 0.8713 1.0000
0.500 0.5448 0.01195 0.00664 -0.1268 0.8658 1.0000
0.750 0.5881 0.01163 0.00626 -0.1296 0.8625 1.0000
1.000 0.6158 0.01161 0.00619 -0.1295 0.8532 1.0000
1.250 0.6566 0.01133 0.00587 -0.1318 0.8484 1.0000
1.750 0.7237 0.01107 0.00553 -0.1336 0.8325 1.0000
2.000 0.7513 0.01104 0.00549 -0.1333 0.8214 1.0000
2.250 0.7805 0.01099 0.00544 -0.1333 0.8106 1.0000
2.500 0.8117 0.01092 0.00535 -0.1337 0.8000 1.0000
2.750 0.8436 0.01085 0.00524 -0.1342 0.7889 1.0000
3.000 0.8721 0.01084 0.00523 -0.1341 0.7760 1.0000
3.250 0.8995 0.01087 0.00524 -0.1337 0.7619 1.0000
3.500 0.9270 0.01090 0.00526 -0.1333 0.7470 1.0000
3.750 0.9529 0.01098 0.00532 -0.1327 0.7305 1.0000
4.000 0.9778 0.01107 0.00543 -0.1319 0.7125 1.0000
4.250 1.0031 0.01118 0.00552 -0.1311 0.6933 1.0000
4.500 1.0287 0.01132 0.00561 -0.1304 0.6736 1.0000
4.750 1.0517 0.01150 0.00580 -0.1292 0.6511 1.0000
5.000 1.0749 0.01172 0.00596 -0.1281 0.6277 1.0000
5.250 1.0971 0.01198 0.00616 -0.1267 0.6021 1.0000
5.500 1.1179 0.01227 0.00640 -0.1252 0.5738 1.0000
5.750 1.1376 0.01263 0.00668 -0.1234 0.5430 1.0000
6.000 1.1561 0.01303 0.00700 -0.1215 0.5095 1.0000
6.250 1.1735 0.01351 0.00737 -0.1194 0.4741 1.0000
6.500 1.1896 0.01406 0.00779 -0.1172 0.4368 1.0000
6.750 1.2043 0.01468 0.00826 -0.1148 0.3984 1.0000
7.000 1.2178 0.01537 0.00882 -0.1122 0.3576 1.0000
7.250 1.2289 0.01619 0.00943 -0.1093 0.3131 1.0000
7.500 1.2381 0.01708 0.01009 -0.1062 0.2689 1.0000
7.750 1.2472 0.01801 0.01081 -0.1031 0.2276 1.0000
8.000 1.2573 0.01898 0.01161 -0.1003 0.1914 1.0000
8.250 1.2674 0.02001 0.01248 -0.0976 0.1590 1.0000
8.500 1.2769 0.02112 0.01344 -0.0949 0.1301 1.0000
8.750 1.2859 0.02231 0.01452 -0.0922 0.1044 1.0000
9.000 1.2934 0.02365 0.01576 -0.0893 0.0825 1.0000
9.250 1.2982 0.02522 0.01722 -0.0861 0.0653 1.0000
9.500 1.3029 0.02684 0.01881 -0.0831 0.0534 1.0000
9.750 1.3092 0.02841 0.02046 -0.0803 0.0463 1.0000
10.000 1.3118 0.03035 0.02238 -0.0773 0.0415 1.0000
10.250 1.3202 0.03196 0.02413 -0.0750 0.0379 1.0000
10.500 1.3292 0.03344 0.02568 -0.0730 0.0344 1.0000
10.750 1.3339 0.03572 0.02799 -0.0707 0.0319 1.0000
11.000 1.3448 0.03774 0.03016 -0.0689 0.0299 1.0000
11.250 1.3553 0.03940 0.03198 -0.0672 0.0278 1.0000
11.500 1.3638 0.04114 0.03381 -0.0656 0.0258 1.0000
11.750 1.3734 0.04345 0.03615 -0.0642 0.0242 1.0000
12.000 1.3895 0.04694 0.03986 -0.0632 0.0229 1.0000
12.250 1.3965 0.04940 0.04260 -0.0615 0.0223 1.0000
12.500 1.4008 0.05220 0.04568 -0.0597 0.0216 1.0000
12.750 1.4014 0.05510 0.04885 -0.0581 0.0208 1.0000
13.000 1.3999 0.05804 0.05201 -0.0566 0.0201 1.0000
13.250 1.3977 0.06103 0.05519 -0.0554 0.0194 1.0000
13.500 1.3932 0.06426 0.05863 -0.0545 0.0188 1.0000
13.750 1.3872 0.06795 0.06250 -0.0538 0.0184 1.0000
14.000 1.3782 0.07227 0.06705 -0.0534 0.0182 1.0000
14.250 1.3661 0.07709 0.07211 -0.0535 0.0181 1.0000
14.500 1.3512 0.08245 0.07773 -0.0542 0.0180 1.0000
14.750 1.3337 0.08840 0.08395 -0.0557 0.0180 1.0000
15.000 1.3140 0.09504 0.09086 -0.0580 0.0181 1.0000
15.250 1.2911 0.10279 0.09889 -0.0615 0.0183 1.0000
15.500 1.2654 0.11172 0.10811 -0.0663 0.0186 1.0000
15.750 1.2239 0.12517 0.12190 -0.0751 0.0194 1.0000
16.000 1.1863 0.13962 0.13662 -0.0853 0.0202 1.0000
16.250 1.1461 0.15707 0.15422 -0.0973 0.0218 1.0000
16.500 1.1223 0.17040 0.16757 -0.1055 0.0229 1.0000
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Polar data table (+)
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