EPPLER 668 AIRFOIL (e668-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 668 AIRFOIL (e668-il) Reynolds number: 200,000 Max Cl/Cd: 77.19 at α=6.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e668-il-200000-n5.txt Download as CSV file: xf-e668-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 668 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
------ -------- --------- --------- -------- -------- --------
-11.500 -0.1388 0.09190 0.08763 -0.1178 0.8992 0.0102
-11.250 -0.1389 0.08693 0.08266 -0.1206 0.8966 0.0101
-10.750 -0.2409 0.05414 0.04979 -0.1379 0.8866 0.0086
-10.500 -0.2664 0.04658 0.04199 -0.1444 0.8823 0.0085
-10.250 -0.2786 0.04327 0.03855 -0.1456 0.8787 0.0084
-10.000 -0.3043 0.03933 0.03435 -0.1449 0.8740 0.0084
-9.750 -0.3234 0.03666 0.03141 -0.1423 0.8697 0.0084
-9.500 -0.3287 0.03467 0.02917 -0.1402 0.8669 0.0084
-9.250 -0.3334 0.03206 0.02618 -0.1379 0.8643 0.0085
-9.000 -0.3324 0.02961 0.02330 -0.1356 0.8614 0.0086
-8.750 -0.3226 0.02797 0.02142 -0.1341 0.8587 0.0088
-8.500 -0.3080 0.02685 0.02016 -0.1331 0.8565 0.0092
-8.250 -0.2915 0.02562 0.01872 -0.1321 0.8546 0.0095
-8.000 -0.2734 0.02439 0.01726 -0.1313 0.8529 0.0098
-7.750 -0.2537 0.02322 0.01586 -0.1305 0.8513 0.0103
-7.500 -0.2330 0.02226 0.01469 -0.1298 0.8496 0.0115
-7.250 -0.2135 0.02175 0.01413 -0.1290 0.8474 0.0123
-7.000 -0.1931 0.02108 0.01330 -0.1281 0.8454 0.0138
-6.750 -0.1727 0.02041 0.01259 -0.1274 0.8434 0.0149
-6.500 -0.1513 0.01974 0.01183 -0.1267 0.8413 0.0164
-6.250 -0.1293 0.01905 0.01105 -0.1260 0.8393 0.0181
-6.000 -0.1055 0.01857 0.01049 -0.1257 0.8377 0.0214
-5.750 -0.0811 0.01809 0.00994 -0.1254 0.8363 0.0254
-5.500 -0.0560 0.01760 0.00940 -0.1253 0.8349 0.0308
-5.250 -0.0341 0.01723 0.00904 -0.1246 0.8327 0.0377
-5.000 -0.0131 0.01691 0.00876 -0.1238 0.8300 0.0472
-4.750 0.0098 0.01660 0.00847 -0.1233 0.8278 0.0612
-4.500 0.0338 0.01625 0.00820 -0.1230 0.8256 0.0821
-4.250 0.0591 0.01592 0.00794 -0.1229 0.8236 0.1114
-4.000 0.0852 0.01555 0.00771 -0.1230 0.8219 0.1517
-3.750 0.1119 0.01518 0.00750 -0.1233 0.8205 0.2056
-3.500 0.1393 0.01480 0.00732 -0.1237 0.8191 0.2704
-3.250 0.1585 0.01462 0.00738 -0.1226 0.8155 0.3343
-3.000 0.1808 0.01436 0.00739 -0.1221 0.8124 0.4084
-2.750 0.2053 0.01406 0.00737 -0.1218 0.8098 0.4906
-2.500 0.2304 0.01380 0.00740 -0.1213 0.8077 0.5763
-2.250 0.2560 0.01370 0.00751 -0.1206 0.8059 0.6471
-2.000 0.2823 0.01375 0.00763 -0.1199 0.8043 0.6941
-1.750 0.3029 0.01398 0.00789 -0.1183 0.8010 0.7251
-1.500 0.3227 0.01423 0.00814 -0.1167 0.7967 0.7482
-1.250 0.3475 0.01437 0.00824 -0.1160 0.7938 0.7666
-1.000 0.3749 0.01443 0.00824 -0.1158 0.7915 0.7816
-0.750 0.4042 0.01446 0.00820 -0.1159 0.7896 0.7949
-0.500 0.4332 0.01447 0.00816 -0.1159 0.7879 0.8057
-0.250 0.4450 0.01480 0.00853 -0.1129 0.7813 0.8154
0.000 0.4707 0.01485 0.00855 -0.1125 0.7779 0.8257
0.250 0.4965 0.01481 0.00849 -0.1118 0.7753 0.8330
0.500 0.5270 0.01473 0.00837 -0.1122 0.7733 0.8418
0.750 0.5411 0.01493 0.00860 -0.1095 0.7675 0.8494
1.000 0.5639 0.01496 0.00863 -0.1086 0.7630 0.8574
1.250 0.5898 0.01483 0.00850 -0.1079 0.7600 0.8636
1.500 0.6213 0.01465 0.00830 -0.1085 0.7577 0.8703
1.750 0.6312 0.01484 0.00855 -0.1051 0.7499 0.8769
2.000 0.6575 0.01469 0.00841 -0.1047 0.7459 0.8825
2.250 0.6901 0.01446 0.00817 -0.1056 0.7431 0.8872
2.500 0.6999 0.01460 0.00837 -0.1022 0.7348 0.8928
2.750 0.7276 0.01439 0.00818 -0.1021 0.7305 0.8972
3.000 0.7501 0.01435 0.00816 -0.1012 0.7243 0.9021
3.500 0.7944 0.01408 0.00797 -0.0989 0.7113 0.9111
4.000 0.8330 0.01394 0.00791 -0.0957 0.6942 0.9215
4.250 0.8597 0.01374 0.00774 -0.0954 0.6858 0.9262
4.500 0.8792 0.01370 0.00774 -0.0938 0.6748 0.9321
4.750 0.9041 0.01355 0.00760 -0.0932 0.6636 0.9376
5.000 0.9315 0.01340 0.00746 -0.0930 0.6505 0.9431
5.250 0.9578 0.01332 0.00738 -0.0927 0.6353 0.9493
5.500 0.9846 0.01329 0.00733 -0.0926 0.6168 0.9569
5.750 1.0136 0.01331 0.00730 -0.0929 0.5954 0.9653
6.250 1.0622 0.01376 0.00756 -0.0923 0.5447 1.0000
6.500 1.0804 0.01420 0.00791 -0.0910 0.5183 1.0000
6.750 1.0962 0.01474 0.00834 -0.0894 0.4918 1.0000
7.000 1.1100 0.01537 0.00885 -0.0875 0.4654 1.0000
7.250 1.1222 0.01607 0.00945 -0.0854 0.4389 1.0000
7.500 1.1336 0.01684 0.01013 -0.0834 0.4124 1.0000
7.750 1.1440 0.01769 0.01087 -0.0813 0.3868 1.0000
8.000 1.1546 0.01857 0.01166 -0.0792 0.3618 1.0000
8.250 1.1649 0.01950 0.01251 -0.0773 0.3375 1.0000
8.500 1.1749 0.02048 0.01341 -0.0753 0.3138 1.0000
8.750 1.1853 0.02148 0.01434 -0.0735 0.2905 1.0000
9.000 1.1946 0.02257 0.01534 -0.0717 0.2672 1.0000
9.250 1.2050 0.02364 0.01635 -0.0700 0.2452 1.0000
9.500 1.2144 0.02480 0.01742 -0.0682 0.2231 1.0000
9.750 1.2238 0.02599 0.01855 -0.0666 0.2020 1.0000
10.000 1.2336 0.02719 0.01969 -0.0650 0.1819 1.0000
10.250 1.2429 0.02845 0.02090 -0.0635 0.1637 1.0000
10.500 1.2520 0.02976 0.02215 -0.0620 0.1467 1.0000
10.750 1.2617 0.03105 0.02342 -0.0606 0.1312 1.0000
11.000 1.2708 0.03241 0.02476 -0.0592 0.1166 1.0000
11.250 1.2798 0.03383 0.02616 -0.0579 0.1036 1.0000
11.500 1.2881 0.03532 0.02764 -0.0566 0.0918 1.0000
11.750 1.2965 0.03685 0.02916 -0.0553 0.0815 1.0000
12.000 1.3052 0.03837 0.03073 -0.0542 0.0726 1.0000
12.250 1.3132 0.04000 0.03240 -0.0530 0.0651 1.0000
12.500 1.3197 0.04181 0.03423 -0.0519 0.0584 1.0000
12.750 1.3273 0.04356 0.03605 -0.0509 0.0526 1.0000
13.000 1.3332 0.04550 0.03805 -0.0499 0.0476 1.0000
13.250 1.3388 0.04754 0.04015 -0.0490 0.0431 1.0000
13.500 1.3438 0.04969 0.04238 -0.0481 0.0395 1.0000
13.750 1.3484 0.05194 0.04472 -0.0474 0.0361 1.0000
14.000 1.3514 0.05442 0.04728 -0.0467 0.0334 1.0000
14.250 1.3554 0.05686 0.04984 -0.0462 0.0307 1.0000
14.500 1.3562 0.05973 0.05278 -0.0458 0.0288 1.0000
14.750 1.3590 0.06246 0.05564 -0.0455 0.0267 1.0000
15.000 1.3608 0.06536 0.05866 -0.0453 0.0248 1.0000
15.500 1.3602 0.07197 0.06551 -0.0455 0.0219 1.0000
15.750 1.3603 0.07534 0.06901 -0.0459 0.0205 1.0000
16.000 1.3575 0.07920 0.07296 -0.0465 0.0194 1.0000
16.250 1.3555 0.08304 0.07696 -0.0471 0.0183 1.0000
16.500 1.3537 0.08695 0.08102 -0.0480 0.0173 1.0000
16.750 1.3507 0.09113 0.08533 -0.0491 0.0163 1.0000
17.000 1.3446 0.09588 0.09019 -0.0506 0.0156 1.0000
17.250 1.3406 0.10040 0.09487 -0.0521 0.0149 1.0000
17.500 1.3362 0.10508 0.09972 -0.0538 0.0141 1.0000
17.750 1.3310 0.10998 0.10477 -0.0558 0.0134 1.0000
18.000 1.3247 0.11517 0.11008 -0.0581 0.0129 1.0000
18.250 1.3167 0.12073 0.11574 -0.0607 0.0125 1.0000
18.500 1.3115 0.12587 0.12107 -0.0632 0.0119 1.0000
18.750 1.3054 0.13123 0.12660 -0.0660 0.0113 1.0000
19.000 1.2988 0.13681 0.13232 -0.0690 0.0108 1.0000
19.250 1.2918 0.14255 0.13819 -0.0724 0.0104 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 668 AIRFOIL (e668-il)