EPPLER 66 AIRFOIL (e66-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: EPPLER 66 AIRFOIL (e66-il) Reynolds number: 500,000 Max Cl/Cd: 128.58 at α=3.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e66-il-500000.txt Download as CSV file: xf-e66-il-500000.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 = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.750 -0.3239 0.06972 0.06762 -0.0707 0.9766 0.0238
-7.500 -0.3367 0.03468 0.03170 -0.1124 0.9614 0.0174
-7.250 -0.3220 0.02317 0.01890 -0.1181 0.9562 0.0134
-7.000 -0.2958 0.02080 0.01617 -0.1191 0.9523 0.0137
-6.750 -0.2639 0.01891 0.01398 -0.1208 0.9503 0.0142
-6.500 -0.2300 0.01741 0.01223 -0.1225 0.9489 0.0150
-6.250 -0.1949 0.01630 0.01091 -0.1243 0.9478 0.0158
-6.000 -0.1600 0.01469 0.00910 -0.1263 0.9469 0.0170
-5.750 -0.1237 0.01394 0.00833 -0.1285 0.9462 0.0191
-5.250 -0.0684 0.01237 0.00657 -0.1288 0.9378 0.0236
-5.000 -0.0325 0.01177 0.00593 -0.1306 0.9365 0.0276
-4.750 0.0046 0.01110 0.00522 -0.1327 0.9354 0.0337
-4.500 0.0420 0.01056 0.00465 -0.1348 0.9342 0.0427
-4.250 0.0793 0.01001 0.00416 -0.1369 0.9330 0.0641
-4.000 0.1063 0.00965 0.00388 -0.1368 0.9278 0.0910
-3.750 0.1384 0.00929 0.00359 -0.1378 0.9242 0.1189
-3.500 0.1738 0.00891 0.00332 -0.1395 0.9216 0.1522
-3.250 0.2100 0.00857 0.00306 -0.1414 0.9193 0.1898
-3.000 0.2373 0.00832 0.00292 -0.1413 0.9133 0.2289
-2.750 0.2693 0.00803 0.00275 -0.1423 0.9086 0.2777
-2.500 0.3038 0.00775 0.00258 -0.1438 0.9050 0.3242
-2.250 0.3307 0.00758 0.00250 -0.1436 0.8981 0.3659
-2.000 0.3621 0.00737 0.00238 -0.1443 0.8927 0.4103
-1.750 0.3916 0.00720 0.00232 -0.1447 0.8865 0.4585
-1.500 0.4203 0.00704 0.00226 -0.1448 0.8795 0.5066
-1.250 0.4498 0.00689 0.00222 -0.1451 0.8729 0.5551
-1.000 0.4775 0.00677 0.00220 -0.1450 0.8650 0.6001
-0.750 0.5057 0.00668 0.00219 -0.1449 0.8573 0.6439
-0.500 0.5336 0.00659 0.00217 -0.1448 0.8491 0.6854
-0.250 0.5598 0.00652 0.00219 -0.1443 0.8401 0.7259
0.000 0.5874 0.00646 0.00219 -0.1440 0.8318 0.7648
0.250 0.6122 0.00642 0.00222 -0.1431 0.8218 0.8012
0.500 0.6366 0.00639 0.00225 -0.1421 0.8119 0.8361
0.750 0.6606 0.00637 0.00226 -0.1410 0.8021 0.8713
1.000 0.6817 0.00633 0.00227 -0.1392 0.7911 0.9073
1.250 0.7026 0.00627 0.00223 -0.1373 0.7795 0.9529
1.500 0.7353 0.00627 0.00220 -0.1383 0.7674 1.0000
1.750 0.7626 0.00636 0.00222 -0.1382 0.7548 1.0000
2.000 0.7895 0.00646 0.00227 -0.1380 0.7415 1.0000
2.250 0.8160 0.00657 0.00232 -0.1377 0.7276 1.0000
2.500 0.8422 0.00669 0.00239 -0.1373 0.7127 1.0000
2.750 0.8679 0.00683 0.00246 -0.1368 0.6967 1.0000
3.000 0.8932 0.00698 0.00256 -0.1362 0.6795 1.0000
3.250 0.9180 0.00715 0.00266 -0.1355 0.6611 1.0000
3.500 0.9425 0.00733 0.00278 -0.1348 0.6416 1.0000
3.750 0.9666 0.00753 0.00292 -0.1340 0.6208 1.0000
4.000 0.9902 0.00776 0.00307 -0.1332 0.5991 1.0000
4.250 1.0132 0.00801 0.00324 -0.1322 0.5747 1.0000
4.500 1.0357 0.00829 0.00343 -0.1311 0.5479 1.0000
4.750 1.0573 0.00861 0.00365 -0.1299 0.5186 1.0000
5.000 1.0782 0.00898 0.00389 -0.1285 0.4869 1.0000
5.250 1.0989 0.00938 0.00416 -0.1272 0.4549 1.0000
5.500 1.1193 0.00980 0.00445 -0.1258 0.4205 1.0000
5.750 1.1394 0.01025 0.00479 -0.1244 0.3870 1.0000
6.000 1.1584 0.01078 0.00515 -0.1229 0.3489 1.0000
6.250 1.1759 0.01143 0.00557 -0.1211 0.3038 1.0000
6.500 1.1931 0.01210 0.00602 -0.1193 0.2601 1.0000
6.750 1.2106 0.01277 0.00649 -0.1176 0.2225 1.0000
7.000 1.2283 0.01340 0.00698 -0.1160 0.1894 1.0000
7.250 1.2450 0.01405 0.00748 -0.1141 0.1596 1.0000
7.500 1.2616 0.01468 0.00800 -0.1122 0.1343 1.0000
7.750 1.2774 0.01536 0.00856 -0.1102 0.1115 1.0000
8.000 1.2930 0.01605 0.00914 -0.1083 0.0904 1.0000
8.250 1.3083 0.01675 0.00978 -0.1062 0.0723 1.0000
8.500 1.3226 0.01752 0.01046 -0.1041 0.0555 1.0000
8.750 1.3350 0.01843 0.01128 -0.1017 0.0398 1.0000
9.000 1.3470 0.01937 0.01219 -0.0992 0.0301 1.0000
9.250 1.3598 0.02025 0.01307 -0.0969 0.0251 1.0000
9.500 1.3717 0.02119 0.01408 -0.0944 0.0220 1.0000
9.750 1.3856 0.02198 0.01492 -0.0924 0.0198 1.0000
10.000 1.3939 0.02318 0.01619 -0.0897 0.0178 1.0000
10.250 1.4013 0.02446 0.01757 -0.0869 0.0166 1.0000
10.500 1.4130 0.02543 0.01864 -0.0848 0.0154 1.0000
10.750 1.4237 0.02649 0.01977 -0.0827 0.0143 1.0000
11.000 1.4302 0.02790 0.02123 -0.0803 0.0132 1.0000
11.250 1.4264 0.03021 0.02367 -0.0768 0.0122 1.0000
11.500 1.4366 0.03142 0.02499 -0.0750 0.0117 1.0000
11.750 1.4438 0.03294 0.02662 -0.0731 0.0111 1.0000
12.000 1.4497 0.03461 0.02840 -0.0711 0.0105 1.0000
12.250 1.4546 0.03642 0.03030 -0.0693 0.0100 1.0000
12.500 1.4580 0.03841 0.03239 -0.0676 0.0096 1.0000
12.750 1.4573 0.04094 0.03503 -0.0657 0.0092 1.0000
13.000 1.4496 0.04448 0.03872 -0.0636 0.0088 1.0000
13.250 1.4486 0.04743 0.04183 -0.0622 0.0086 1.0000
13.500 1.4514 0.04987 0.04442 -0.0612 0.0084 1.0000
13.750 1.4523 0.05265 0.04736 -0.0603 0.0082 1.0000
14.000 1.4519 0.05569 0.05056 -0.0596 0.0080 1.0000
14.250 1.4503 0.05900 0.05403 -0.0592 0.0078 1.0000
14.500 1.4476 0.06253 0.05773 -0.0590 0.0076 1.0000
14.750 1.4440 0.06633 0.06168 -0.0591 0.0074 1.0000
15.000 1.4403 0.07025 0.06575 -0.0596 0.0072 1.0000
15.250 1.4355 0.07445 0.07009 -0.0604 0.0071 1.0000
15.500 1.4296 0.07902 0.07480 -0.0615 0.0069 1.0000
15.750 1.4234 0.08380 0.07972 -0.0630 0.0068 1.0000
16.000 1.4160 0.08897 0.08502 -0.0649 0.0067 1.0000
16.250 1.4071 0.09457 0.09076 -0.0671 0.0066 1.0000
16.500 1.3973 0.10053 0.09689 -0.0697 0.0065 1.0000
16.750 1.3865 0.10687 0.10338 -0.0727 0.0064 1.0000
17.000 1.3745 0.11370 0.11036 -0.0762 0.0064 1.0000
17.250 1.3613 0.12098 0.11781 -0.0802 0.0063 1.0000
17.500 1.3473 0.12867 0.12567 -0.0847 0.0063 1.0000
17.750 1.3333 0.13667 0.13384 -0.0896 0.0063 1.0000
18.000 1.3191 0.14509 0.14243 -0.0950 0.0063 1.0000
18.250 1.3031 0.15425 0.15177 -0.1010 0.0064 1.0000
18.500 1.2835 0.16488 0.16258 -0.1082 0.0065 1.0000
18.750 1.2387 0.18438 0.18243 -0.1213 0.0068 1.0000
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Polar data table (+)
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