EPPLER 682 AIRFOIL (e682-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 682 AIRFOIL (e682-il) Reynolds number: 100,000 Max Cl/Cd: 37.85 at α=9° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e682-il-100000-n5.txt Download as CSV file: xf-e682-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 682 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.000 -0.2621 0.10556 0.10042 -0.0906 0.9301 0.0329
-11.750 -0.2537 0.10188 0.09672 -0.0925 0.9257 0.0325
-11.500 -0.2475 0.09807 0.09289 -0.0945 0.9210 0.0318
-11.250 -0.2439 0.09372 0.08853 -0.0972 0.9168 0.0310
-11.000 -0.2447 0.08829 0.08308 -0.1007 0.9129 0.0301
-10.750 -0.2511 0.08229 0.07710 -0.1045 0.9075 0.0293
-10.500 -0.3145 0.06546 0.05996 -0.1170 0.8997 0.0265
-10.250 -0.3266 0.06190 0.05627 -0.1174 0.8950 0.0264
-10.000 -0.3392 0.05884 0.05309 -0.1168 0.8900 0.0262
-9.750 -0.3515 0.05605 0.05014 -0.1153 0.8855 0.0261
-9.500 -0.3635 0.05346 0.04736 -0.1129 0.8817 0.0259
-9.250 -0.3754 0.05142 0.04515 -0.1097 0.8771 0.0258
-9.000 -0.3827 0.04907 0.04256 -0.1067 0.8729 0.0257
-8.750 -0.3859 0.04667 0.03988 -0.1039 0.8694 0.0255
-8.500 -0.3864 0.04412 0.03697 -0.1011 0.8669 0.0255
-8.250 -0.3861 0.04216 0.03472 -0.0980 0.8634 0.0254
-8.000 -0.3822 0.04019 0.03243 -0.0952 0.8601 0.0254
-7.750 -0.3745 0.03819 0.03007 -0.0927 0.8572 0.0254
-7.500 -0.3620 0.03623 0.02773 -0.0907 0.8547 0.0255
-7.250 -0.3448 0.03438 0.02551 -0.0893 0.8527 0.0257
-7.000 -0.3252 0.03278 0.02357 -0.0882 0.8510 0.0260
-6.750 -0.3112 0.03174 0.02229 -0.0862 0.8480 0.0268
-6.500 -0.2949 0.03088 0.02139 -0.0848 0.8450 0.0280
-6.250 -0.2750 0.03003 0.02043 -0.0838 0.8424 0.0294
-6.000 -0.2516 0.02903 0.01928 -0.0832 0.8405 0.0304
-5.750 -0.2269 0.02804 0.01816 -0.0828 0.8388 0.0312
-5.500 -0.2022 0.02715 0.01714 -0.0823 0.8372 0.0322
-5.250 -0.1848 0.02657 0.01648 -0.0807 0.8346 0.0333
-5.000 -0.1780 0.02634 0.01627 -0.0774 0.8305 0.0345
-4.750 -0.1636 0.02598 0.01587 -0.0753 0.8273 0.0364
-4.500 -0.1449 0.02561 0.01540 -0.0739 0.8248 0.0402
-4.250 -0.1247 0.02511 0.01487 -0.0728 0.8227 0.0457
-4.000 -0.1020 0.02456 0.01430 -0.0719 0.8210 0.0556
-3.750 -0.1032 0.02468 0.01452 -0.0672 0.8151 0.0700
-3.500 -0.0902 0.02422 0.01437 -0.0649 0.8115 0.1285
-3.250 -0.0757 0.02320 0.01423 -0.0631 0.8089 0.2985
-3.000 -0.0346 0.02385 0.01696 -0.0606 0.8087 0.7713
-2.750 -0.0309 0.02430 0.01725 -0.0555 0.8059 0.8128
-2.500 -0.0514 0.02505 0.01799 -0.0469 0.7977 0.8310
-2.250 -0.0169 0.02584 0.01858 -0.0468 0.7954 0.8524
-2.000 0.0247 0.02649 0.01904 -0.0481 0.7939 0.8751
-1.750 0.1054 0.02701 0.01929 -0.0571 0.7945 0.8943
-1.500 0.1555 0.02704 0.01913 -0.0613 0.7933 0.9032
-1.250 0.1965 0.02693 0.01887 -0.0640 0.7917 0.9091
-1.000 0.2386 0.02678 0.01857 -0.0670 0.7903 0.9138
-0.750 0.2373 0.02741 0.01919 -0.0624 0.7823 0.9229
-0.500 0.2715 0.02735 0.01903 -0.0641 0.7794 0.9269
-0.250 0.3003 0.02723 0.01882 -0.0645 0.7769 0.9316
0.000 0.3316 0.02705 0.01855 -0.0654 0.7749 0.9352
0.250 0.3414 0.02761 0.01911 -0.0632 0.7667 0.9406
0.500 0.3666 0.02756 0.01900 -0.0630 0.7631 0.9444
0.750 0.3922 0.02738 0.01877 -0.0627 0.7605 0.9477
1.000 0.4289 0.02716 0.01849 -0.0646 0.7584 0.9492
1.500 0.4631 0.02757 0.01890 -0.0620 0.7458 0.9559
1.750 0.4955 0.02728 0.01859 -0.0629 0.7437 0.9576
2.000 0.4891 0.02798 0.01933 -0.0577 0.7328 0.9624
2.250 0.5230 0.02769 0.01903 -0.0590 0.7301 0.9635
2.500 0.5542 0.02745 0.01879 -0.0598 0.7271 0.9647
2.750 0.5540 0.02800 0.01938 -0.0556 0.7165 0.9687
3.000 0.5867 0.02760 0.01900 -0.0564 0.7140 0.9695
3.500 0.6126 0.02774 0.01920 -0.0518 0.6995 0.9742
4.000 0.6459 0.02774 0.01931 -0.0487 0.6849 0.9783
4.500 0.6611 0.02812 0.01978 -0.0427 0.6665 0.9831
4.750 0.6767 0.02819 0.01992 -0.0411 0.6577 0.9852
5.000 0.7107 0.02744 0.01924 -0.0417 0.6550 0.9857
5.250 0.7167 0.02790 0.01977 -0.0390 0.6425 0.9880
5.500 0.7267 0.02823 0.02017 -0.0369 0.6312 0.9902
5.750 0.7586 0.02745 0.01946 -0.0370 0.6271 0.9911
6.000 0.7710 0.02774 0.01986 -0.0352 0.6153 0.9938
6.250 0.7896 0.02774 0.01994 -0.0341 0.6051 0.9958
6.500 0.8208 0.02698 0.01926 -0.0341 0.5987 0.9968
6.750 0.8386 0.02700 0.01939 -0.0328 0.5865 0.9989
7.000 0.8575 0.02677 0.01923 -0.0313 0.5747 1.0000
7.250 0.8778 0.02627 0.01878 -0.0295 0.5640 1.0000
7.500 0.8999 0.02570 0.01825 -0.0279 0.5514 1.0000
7.750 0.9179 0.02535 0.01795 -0.0259 0.5373 1.0000
8.000 0.9310 0.02521 0.01783 -0.0233 0.5218 1.0000
8.250 0.9403 0.02523 0.01786 -0.0202 0.5050 1.0000
8.500 0.9489 0.02530 0.01795 -0.0171 0.4878 1.0000
8.750 0.9574 0.02542 0.01807 -0.0141 0.4697 1.0000
9.000 0.9681 0.02558 0.01822 -0.0115 0.4500 1.0000
9.250 0.9787 0.02591 0.01853 -0.0091 0.4292 1.0000
9.500 0.9877 0.02647 0.01907 -0.0069 0.4069 1.0000
9.750 0.9989 0.02708 0.01963 -0.0051 0.3846 1.0000
10.000 1.0076 0.02791 0.02045 -0.0033 0.3608 1.0000
10.250 1.0162 0.02884 0.02134 -0.0016 0.3370 1.0000
10.500 1.0239 0.02991 0.02236 0.0001 0.3129 1.0000
10.750 1.0308 0.03111 0.02351 0.0016 0.2889 1.0000
11.000 1.0365 0.03245 0.02478 0.0032 0.2652 1.0000
11.250 1.0416 0.03391 0.02619 0.0046 0.2405 1.0000
11.500 1.0447 0.03557 0.02777 0.0060 0.2150 1.0000
11.750 1.0461 0.03745 0.02953 0.0073 0.1878 1.0000
12.000 1.0475 0.03942 0.03140 0.0085 0.1635 1.0000
12.250 1.0489 0.04153 0.03339 0.0095 0.1378 1.0000
12.500 1.0504 0.04371 0.03546 0.0104 0.1135 1.0000
12.750 1.0526 0.04593 0.03762 0.0112 0.0949 1.0000
13.000 1.0556 0.04817 0.03984 0.0118 0.0785 1.0000
13.250 1.0565 0.05069 0.04232 0.0123 0.0635 1.0000
13.500 1.0558 0.05346 0.04505 0.0128 0.0512 1.0000
13.750 1.0553 0.05630 0.04793 0.0131 0.0427 1.0000
14.000 1.0532 0.05946 0.05114 0.0133 0.0363 1.0000
14.250 1.0496 0.06287 0.05461 0.0133 0.0329 1.0000
14.500 1.0486 0.06611 0.05799 0.0132 0.0304 1.0000
14.750 1.0469 0.06953 0.06155 0.0129 0.0284 1.0000
15.000 1.0439 0.07320 0.06534 0.0123 0.0268 1.0000
15.250 1.0406 0.07698 0.06922 0.0116 0.0259 1.0000
15.500 1.0396 0.08054 0.07292 0.0109 0.0247 1.0000
15.750 1.0395 0.08404 0.07659 0.0101 0.0237 1.0000
16.000 1.0392 0.08764 0.08032 0.0092 0.0227 1.0000
16.250 1.0386 0.09136 0.08417 0.0080 0.0218 1.0000
16.500 1.0369 0.09528 0.08818 0.0066 0.0210 1.0000
16.750 1.0353 0.09916 0.09213 0.0053 0.0204 1.0000
17.000 1.0367 0.10268 0.09583 0.0042 0.0198 1.0000
17.250 1.0387 0.10612 0.09945 0.0031 0.0194 1.0000
17.500 1.0400 0.10977 0.10326 0.0018 0.0190 1.0000
17.750 1.0399 0.11368 0.10736 0.0002 0.0186 1.0000
18.000 1.0385 0.11795 0.11180 -0.0016 0.0183 1.0000
18.250 1.0362 0.12244 0.11647 -0.0037 0.0181 1.0000
18.500 1.0320 0.12742 0.12163 -0.0062 0.0179 1.0000
18.750 1.0262 0.13286 0.12726 -0.0092 0.0177 1.0000
19.000 1.0183 0.13891 0.13351 -0.0126 0.0177 1.0000
19.250 1.0079 0.14570 0.14050 -0.0167 0.0177 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 682 AIRFOIL (e682-il)