EPPLER 67 AIRFOIL (e67-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: EPPLER 67 AIRFOIL (e67-il) Reynolds number: 200,000 Max Cl/Cd: 86.82 at α=5.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e67-il-200000.txt Download as CSV file: xf-e67-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 67 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
------ -------- --------- --------- -------- -------- --------
-8.750 -0.4163 0.09843 0.09525 -0.0338 1.0000 0.0589
-8.500 -0.4288 0.09592 0.09280 -0.0329 1.0000 0.0604
-8.250 -0.4479 0.09321 0.09017 -0.0326 1.0000 0.0621
-7.250 -0.4693 0.04807 0.04423 -0.0802 0.9719 0.0430
-7.000 -0.4646 0.03557 0.03024 -0.0840 0.9641 0.0315
-6.750 -0.4327 0.03297 0.02720 -0.0863 0.9608 0.0311
-6.500 -0.3998 0.02929 0.02296 -0.0890 0.9588 0.0311
-6.250 -0.3784 0.02616 0.01941 -0.0890 0.9535 0.0317
-6.000 -0.3476 0.02406 0.01710 -0.0903 0.9501 0.0330
-5.750 -0.3123 0.02277 0.01570 -0.0921 0.9475 0.0354
-5.500 -0.2734 0.02200 0.01475 -0.0945 0.9454 0.0401
-5.250 -0.2514 0.02061 0.01331 -0.0938 0.9392 0.0441
-5.000 -0.2174 0.01956 0.01219 -0.0952 0.9357 0.0535
-4.750 -0.1797 0.01865 0.01133 -0.0975 0.9332 0.0742
-4.500 -0.1398 0.01802 0.01075 -0.1003 0.9314 0.1014
-4.250 -0.1191 0.01762 0.01041 -0.0992 0.9235 0.1245
-4.000 -0.0822 0.01709 0.00995 -0.1013 0.9204 0.1602
-3.750 -0.0424 0.01652 0.00953 -0.1040 0.9184 0.2051
-3.500 -0.0182 0.01621 0.00937 -0.1036 0.9116 0.2449
-3.250 0.0166 0.01578 0.00910 -0.1053 0.9077 0.2980
-3.000 0.0555 0.01533 0.00883 -0.1076 0.9053 0.3593
-2.750 0.0963 0.01490 0.00858 -0.1102 0.9036 0.4233
-2.500 0.1171 0.01481 0.00860 -0.1089 0.8952 0.4708
-2.250 0.1550 0.01445 0.00840 -0.1108 0.8925 0.5273
-2.000 0.1955 0.01408 0.00815 -0.1131 0.8906 0.5800
-1.750 0.2255 0.01391 0.00808 -0.1133 0.8855 0.6241
-1.500 0.2562 0.01373 0.00799 -0.1136 0.8801 0.6661
-1.250 0.2941 0.01342 0.00777 -0.1151 0.8775 0.7075
-1.000 0.3325 0.01310 0.00751 -0.1166 0.8754 0.7452
-0.750 0.3516 0.01309 0.00758 -0.1145 0.8666 0.7773
-0.500 0.3854 0.01281 0.00735 -0.1150 0.8631 0.8113
-0.250 0.4204 0.01250 0.00707 -0.1157 0.8605 0.8436
0.000 0.4339 0.01249 0.00712 -0.1122 0.8506 0.8753
0.250 0.4651 0.01216 0.00680 -0.1120 0.8470 0.9081
0.500 0.4888 0.01206 0.00673 -0.1107 0.8384 0.9470
0.750 0.5426 0.01175 0.00638 -0.1157 0.8346 0.9801
1.000 0.5823 0.01161 0.00619 -0.1183 0.8278 1.0000
1.250 0.6142 0.01149 0.00602 -0.1192 0.8203 1.0000
1.500 0.6430 0.01147 0.00595 -0.1195 0.8118 1.0000
1.750 0.6772 0.01134 0.00576 -0.1207 0.8046 1.0000
2.000 0.7036 0.01137 0.00577 -0.1204 0.7945 1.0000
2.250 0.7356 0.01131 0.00566 -0.1211 0.7862 1.0000
2.500 0.7649 0.01129 0.00560 -0.1212 0.7763 1.0000
2.750 0.7913 0.01132 0.00562 -0.1208 0.7651 1.0000
3.000 0.8194 0.01133 0.00563 -0.1207 0.7540 1.0000
3.250 0.8484 0.01134 0.00560 -0.1208 0.7429 1.0000
3.500 0.8774 0.01135 0.00558 -0.1208 0.7312 1.0000
3.750 0.9030 0.01142 0.00566 -0.1202 0.7177 1.0000
4.000 0.9286 0.01150 0.00572 -0.1196 0.7036 1.0000
4.250 0.9543 0.01158 0.00579 -0.1190 0.6890 1.0000
4.500 0.9797 0.01169 0.00588 -0.1183 0.6734 1.0000
4.750 1.0030 0.01183 0.00604 -0.1173 0.6564 1.0000
5.000 1.0261 0.01198 0.00619 -0.1162 0.6382 1.0000
5.250 1.0495 0.01215 0.00634 -0.1152 0.6193 1.0000
5.500 1.0721 0.01236 0.00651 -0.1140 0.5991 1.0000
5.750 1.0931 0.01259 0.00672 -0.1126 0.5767 1.0000
6.000 1.1135 0.01286 0.00698 -0.1110 0.5533 1.0000
6.250 1.1329 0.01318 0.00723 -0.1093 0.5279 1.0000
6.500 1.1507 0.01355 0.00753 -0.1073 0.5001 1.0000
6.750 1.1669 0.01397 0.00788 -0.1051 0.4697 1.0000
7.000 1.1819 0.01445 0.00828 -0.1027 0.4377 1.0000
7.250 1.1951 0.01501 0.00874 -0.1000 0.4037 1.0000
7.500 1.2059 0.01565 0.00925 -0.0970 0.3693 1.0000
7.750 1.2143 0.01634 0.00980 -0.0936 0.3326 1.0000
8.000 1.2212 0.01718 0.01046 -0.0901 0.2964 1.0000
8.250 1.2279 0.01810 0.01120 -0.0867 0.2581 1.0000
8.500 1.2343 0.01911 0.01202 -0.0835 0.2223 1.0000
8.750 1.2414 0.02017 0.01292 -0.0805 0.1891 1.0000
9.000 1.2487 0.02129 0.01392 -0.0777 0.1604 1.0000
9.250 1.2561 0.02246 0.01498 -0.0750 0.1352 1.0000
9.500 1.2624 0.02375 0.01616 -0.0723 0.1133 1.0000
9.750 1.2684 0.02513 0.01745 -0.0696 0.0929 1.0000
10.000 1.2714 0.02676 0.01897 -0.0668 0.0762 1.0000
10.250 1.2751 0.02844 0.02063 -0.0641 0.0618 1.0000
10.500 1.2752 0.03044 0.02260 -0.0612 0.0523 1.0000
10.750 1.2807 0.03212 0.02435 -0.0590 0.0455 1.0000
11.000 1.2809 0.03434 0.02659 -0.0565 0.0411 1.0000
11.250 1.2884 0.03596 0.02834 -0.0548 0.0369 1.0000
11.500 1.2904 0.03817 0.03056 -0.0529 0.0341 1.0000
11.750 1.2948 0.04049 0.03298 -0.0510 0.0319 1.0000
12.000 1.3030 0.04232 0.03496 -0.0496 0.0297 1.0000
12.250 1.3096 0.04425 0.03697 -0.0484 0.0276 1.0000
12.500 1.3143 0.04657 0.03931 -0.0471 0.0257 1.0000
12.750 1.3246 0.04933 0.04220 -0.0458 0.0243 1.0000
13.000 1.3326 0.05159 0.04467 -0.0446 0.0235 1.0000
13.250 1.3397 0.05413 0.04742 -0.0435 0.0226 1.0000
13.500 1.3453 0.05690 0.05040 -0.0425 0.0218 1.0000
13.750 1.3478 0.05980 0.05351 -0.0417 0.0211 1.0000
14.000 1.3477 0.06288 0.05678 -0.0410 0.0204 1.0000
14.250 1.3463 0.06602 0.06005 -0.0406 0.0197 1.0000
14.500 1.3449 0.06938 0.06352 -0.0404 0.0191 1.0000
14.750 1.3404 0.07373 0.06804 -0.0403 0.0185 1.0000
15.000 1.3283 0.07924 0.07382 -0.0406 0.0182 1.0000
15.250 1.3144 0.08477 0.07962 -0.0415 0.0181 1.0000
15.500 1.2979 0.09075 0.08586 -0.0430 0.0181 1.0000
15.750 1.2806 0.09703 0.09239 -0.0451 0.0181 1.0000
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Polar data table (+)
Polar graphs
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