EPPLER 326 AIRFOIL (e326-il) Xfoil prediction polar at RE=50,000 Ncrit=5
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Airfoil: EPPLER 326 AIRFOIL (e326-il) Reynolds number: 50,000 Max Cl/Cd: 26.93 at α=5.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e326-il-50000-n5.txt Download as CSV file: xf-e326-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 326 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.4959 0.10033 0.09362 -0.0146 1.0000 0.0459
-9.750 -0.5028 0.09444 0.08777 -0.0186 1.0000 0.0441
-9.500 -0.5429 0.08647 0.07966 -0.0250 1.0000 0.0408
-9.250 -0.5423 0.08286 0.07609 -0.0248 1.0000 0.0402
-9.000 -0.5486 0.07942 0.07265 -0.0243 1.0000 0.0397
-8.750 -0.5583 0.07609 0.06930 -0.0230 1.0000 0.0393
-8.500 -0.5682 0.07296 0.06613 -0.0211 1.0000 0.0388
-8.250 -0.5765 0.06961 0.06269 -0.0191 1.0000 0.0383
-8.000 -0.5823 0.06633 0.05928 -0.0170 1.0000 0.0377
-7.750 -0.5860 0.06309 0.05589 -0.0146 1.0000 0.0372
-7.500 -0.5883 0.05979 0.05239 -0.0120 1.0000 0.0367
-7.250 -0.5878 0.05667 0.04903 -0.0093 1.0000 0.0361
-7.000 -0.5856 0.05355 0.04565 -0.0065 1.0000 0.0357
-6.750 -0.5814 0.05054 0.04235 -0.0035 1.0000 0.0354
-6.500 -0.5747 0.04772 0.03922 -0.0006 1.0000 0.0352
-6.250 -0.5659 0.04506 0.03625 0.0022 1.0000 0.0350
-6.000 -0.5553 0.04269 0.03359 0.0049 1.0000 0.0353
-5.750 -0.5437 0.04059 0.03125 0.0075 1.0000 0.0364
-5.500 -0.5311 0.03867 0.02907 0.0099 1.0000 0.0376
-5.250 -0.5184 0.03685 0.02695 0.0125 1.0000 0.0392
-5.000 -0.4910 0.03479 0.02443 0.0126 0.9908 0.0406
-4.750 -0.4414 0.03207 0.02153 0.0086 0.9702 0.0427
-4.500 -0.3863 0.02991 0.01918 0.0042 0.9494 0.0480
-4.250 -0.3205 0.02788 0.01702 -0.0018 0.9318 0.0565
-4.000 -0.2656 0.02621 0.01519 -0.0062 0.9085 0.0722
-3.750 -0.2235 0.02464 0.01363 -0.0087 0.8819 0.1027
-3.500 -0.1736 0.02298 0.01550 -0.0071 0.8658 0.7910
-3.250 -0.0138 0.02618 0.01749 -0.0222 0.8411 0.9475
-3.000 0.0242 0.02587 0.01668 -0.0243 0.8070 0.9568
-2.750 0.0596 0.02548 0.01587 -0.0260 0.7782 0.9637
-2.500 0.0892 0.02528 0.01529 -0.0267 0.7541 0.9713
-2.250 0.1221 0.02490 0.01459 -0.0281 0.7318 0.9769
-2.000 0.1513 0.02468 0.01408 -0.0287 0.7124 0.9828
-1.750 0.1827 0.02436 0.01348 -0.0298 0.6945 0.9877
-1.500 0.2128 0.02412 0.01300 -0.0308 0.6777 0.9927
-1.250 0.2435 0.02386 0.01253 -0.0319 0.6621 0.9972
-1.000 0.2707 0.02372 0.01221 -0.0323 0.6476 1.0000
-0.750 0.2924 0.02372 0.01208 -0.0316 0.6341 1.0000
-0.500 0.3142 0.02375 0.01199 -0.0309 0.6212 1.0000
-0.250 0.3359 0.02379 0.01191 -0.0302 0.6094 1.0000
0.000 0.3577 0.02384 0.01183 -0.0294 0.5989 1.0000
0.250 0.3797 0.02393 0.01186 -0.0288 0.5876 1.0000
0.500 0.4016 0.02403 0.01190 -0.0281 0.5769 1.0000
0.750 0.4233 0.02414 0.01189 -0.0273 0.5680 1.0000
1.000 0.4450 0.02430 0.01206 -0.0266 0.5577 1.0000
1.250 0.4668 0.02447 0.01217 -0.0259 0.5485 1.0000
1.500 0.4888 0.02464 0.01231 -0.0251 0.5397 1.0000
1.750 0.5108 0.02487 0.01255 -0.0245 0.5306 1.0000
2.000 0.5326 0.02506 0.01268 -0.0237 0.5230 1.0000
2.250 0.5542 0.02535 0.01303 -0.0231 0.5139 1.0000
2.500 0.5758 0.02557 0.01322 -0.0222 0.5071 1.0000
2.750 0.5968 0.02593 0.01366 -0.0215 0.4982 1.0000
3.000 0.6183 0.02615 0.01384 -0.0205 0.4915 1.0000
3.250 0.6388 0.02659 0.01439 -0.0198 0.4831 1.0000
3.500 0.6600 0.02685 0.01467 -0.0188 0.4766 1.0000
3.750 0.6799 0.02733 0.01527 -0.0180 0.4683 1.0000
4.000 0.7007 0.02765 0.01561 -0.0170 0.4619 1.0000
4.250 0.7202 0.02817 0.01625 -0.0161 0.4543 1.0000
4.500 0.7402 0.02855 0.01672 -0.0151 0.4474 1.0000
4.750 0.7594 0.02907 0.01735 -0.0140 0.4406 1.0000
5.000 0.7781 0.02960 0.01799 -0.0129 0.4333 1.0000
5.250 0.7975 0.03002 0.01850 -0.0117 0.4271 1.0000
5.500 0.8141 0.03075 0.01941 -0.0105 0.4193 1.0000
5.750 0.8347 0.03100 0.01967 -0.0092 0.4139 1.0000
6.000 0.8481 0.03199 0.02093 -0.0077 0.4053 1.0000
6.250 0.8677 0.03231 0.02131 -0.0064 0.3995 1.0000
6.500 0.8797 0.03333 0.02257 -0.0047 0.3913 1.0000
6.750 0.8981 0.03369 0.02301 -0.0032 0.3850 1.0000
7.000 0.9088 0.03473 0.02426 -0.0013 0.3771 1.0000
7.250 0.9255 0.03515 0.02481 0.0004 0.3704 1.0000
7.500 0.9347 0.03619 0.02605 0.0026 0.3627 1.0000
7.750 0.9493 0.03667 0.02665 0.0045 0.3556 1.0000
8.000 0.9563 0.03772 0.02790 0.0070 0.3478 1.0000
8.250 0.9691 0.03822 0.02851 0.0091 0.3406 1.0000
8.500 0.9723 0.03937 0.02988 0.0120 0.3328 1.0000
8.750 0.9840 0.03979 0.03041 0.0144 0.3254 1.0000
9.000 0.9810 0.04113 0.03194 0.0179 0.3177 1.0000
9.250 0.9919 0.04144 0.03235 0.0205 0.3102 1.0000
9.500 0.9803 0.04310 0.03417 0.0247 0.3030 1.0000
9.750 0.9888 0.04338 0.03455 0.0276 0.2954 1.0000
10.000 0.9676 0.04542 0.03669 0.0324 0.2888 1.0000
10.250 0.9616 0.04653 0.03789 0.0361 0.2818 1.0000
10.500 0.9432 0.04822 0.03960 0.0406 0.2758 1.0000
10.750 0.8705 0.05527 0.04654 0.0436 0.2694 1.0000
11.000 0.8998 0.05389 0.04530 0.0456 0.2622 1.0000
11.250 0.8443 0.06216 0.05348 0.0442 0.2526 1.0000
11.500 0.7972 0.07118 0.06242 0.0413 0.2412 1.0000
11.750 0.8488 0.06602 0.05753 0.0451 0.2365 1.0000
12.000 0.7954 0.07662 0.06799 0.0411 0.2245 1.0000
12.250 0.7725 0.08317 0.07452 0.0390 0.2146 1.0000
12.500 0.7999 0.08134 0.07289 0.0410 0.2089 1.0000
13.000 0.8118 0.08489 0.07662 0.0413 0.1937 1.0000
13.250 0.7647 0.09645 0.08806 0.0362 0.1828 1.0000
13.750 0.7641 0.10291 0.09465 0.0347 0.1681 1.0000
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Polar data table (+)
Polar graphs
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