GOE 210 (DAIMLER) AIRFOIL (goe210-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 210 (DAIMLER) AIRFOIL (goe210-il) Reynolds number: 100,000 Max Cl/Cd: 62.04 at α=4° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe210-il-100000-n5.txt Download as CSV file: xf-goe210-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 210 (DAIMLER) 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
------ -------- --------- --------- -------- -------- --------
-7.750 -0.3304 0.09710 0.09235 -0.0254 1.0000 0.0323
-7.500 -0.3324 0.09521 0.09056 -0.0260 1.0000 0.0329
-7.250 -0.3329 0.09330 0.08875 -0.0279 1.0000 0.0334
-7.000 -0.3304 0.09111 0.08664 -0.0309 1.0000 0.0337
-6.750 -0.3256 0.08868 0.08426 -0.0338 1.0000 0.0339
-6.500 -0.3197 0.08607 0.08169 -0.0362 1.0000 0.0340
-6.250 -0.3135 0.08338 0.07902 -0.0381 1.0000 0.0341
-6.000 -0.3147 0.07943 0.07517 -0.0366 1.0000 0.0345
-5.750 -0.3063 0.07519 0.07099 -0.0343 0.9967 0.0360
-5.500 -0.2778 0.07074 0.06647 -0.0398 0.9903 0.0375
-5.250 -0.2444 0.06616 0.06181 -0.0470 0.9833 0.0393
-5.000 -0.2049 0.06134 0.05686 -0.0556 0.9770 0.0420
-4.500 -0.1200 0.05129 0.04634 -0.0724 0.9617 0.0467
-4.250 -0.0942 0.04836 0.04341 -0.0745 0.9556 0.0530
-3.750 -0.0044 0.03723 0.03134 -0.0857 0.9429 0.0331
-3.500 0.0366 0.03313 0.02678 -0.0895 0.9386 0.0332
-3.250 0.0659 0.03065 0.02410 -0.0913 0.9298 0.0373
-3.000 0.1045 0.02755 0.02053 -0.0938 0.9250 0.0381
-2.750 0.1368 0.02525 0.01777 -0.0948 0.9167 0.0401
-2.500 0.1739 0.02317 0.01509 -0.0962 0.9115 0.0442
-2.250 0.2060 0.02153 0.01295 -0.0967 0.9036 0.0455
-2.000 0.2387 0.02027 0.01161 -0.0980 0.8976 0.0510
-1.750 0.2697 0.01914 0.01022 -0.0983 0.8901 0.0530
-1.500 0.3023 0.01814 0.00899 -0.0989 0.8839 0.0550
-1.250 0.3323 0.01751 0.00820 -0.0990 0.8762 0.0586
-1.000 0.3633 0.01681 0.00752 -0.0996 0.8698 0.0628
-0.750 0.3919 0.01629 0.00695 -0.0995 0.8617 0.0646
-0.500 0.4227 0.01584 0.00641 -0.0998 0.8552 0.0671
-0.250 0.4506 0.01557 0.00605 -0.0996 0.8463 0.0707
0.000 0.4803 0.01527 0.00568 -0.0996 0.8383 0.0782
0.250 0.5098 0.01494 0.00532 -0.0994 0.8272 0.1027
0.500 0.5354 0.01267 0.00514 -0.0985 0.8137 1.0000
0.750 0.5627 0.01269 0.00490 -0.0977 0.7981 1.0000
1.000 0.5893 0.01275 0.00476 -0.0970 0.7824 1.0000
1.250 0.6154 0.01284 0.00469 -0.0962 0.7664 1.0000
1.500 0.6412 0.01293 0.00466 -0.0954 0.7495 1.0000
1.750 0.6671 0.01304 0.00465 -0.0947 0.7332 1.0000
2.000 0.6932 0.01315 0.00469 -0.0940 0.7184 1.0000
2.250 0.7194 0.01328 0.00477 -0.0935 0.7042 1.0000
2.500 0.7453 0.01341 0.00485 -0.0929 0.6885 1.0000
2.750 0.7711 0.01354 0.00494 -0.0922 0.6715 1.0000
3.000 0.7967 0.01368 0.00506 -0.0915 0.6533 1.0000
3.250 0.8220 0.01383 0.00520 -0.0908 0.6327 1.0000
3.500 0.8471 0.01399 0.00533 -0.0901 0.6091 1.0000
3.750 0.8717 0.01418 0.00546 -0.0892 0.5793 1.0000
4.000 0.8952 0.01443 0.00560 -0.0881 0.5362 1.0000
4.250 0.9170 0.01486 0.00574 -0.0867 0.4778 1.0000
4.500 0.9373 0.01551 0.00603 -0.0853 0.4157 1.0000
4.750 0.9576 0.01624 0.00649 -0.0841 0.3651 1.0000
5.000 0.9789 0.01692 0.00699 -0.0831 0.3291 1.0000
5.250 1.0008 0.01756 0.00752 -0.0822 0.3020 1.0000
5.500 1.0232 0.01815 0.00807 -0.0814 0.2798 1.0000
5.750 1.0453 0.01877 0.00866 -0.0805 0.2606 1.0000
6.000 1.0674 0.01937 0.00924 -0.0797 0.2445 1.0000
6.250 1.0900 0.01993 0.00985 -0.0790 0.2322 1.0000
6.500 1.1127 0.02048 0.01049 -0.0782 0.2218 1.0000
6.750 1.1349 0.02108 0.01119 -0.0774 0.2118 1.0000
7.250 1.1780 0.02233 0.01259 -0.0757 0.1893 1.0000
7.500 1.1984 0.02300 0.01328 -0.0748 0.1700 1.0000
7.750 1.2189 0.02365 0.01400 -0.0740 0.1498 1.0000
8.000 1.2396 0.02430 0.01472 -0.0731 0.1294 1.0000
8.250 1.2599 0.02502 0.01553 -0.0721 0.1013 1.0000
8.500 1.2757 0.02622 0.01652 -0.0707 0.0760 1.0000
8.750 1.2884 0.02785 0.01799 -0.0690 0.0561 1.0000
9.250 1.3183 0.03051 0.02078 -0.0657 0.0247 1.0000
9.500 1.3276 0.03236 0.02259 -0.0636 0.0202 1.0000
9.750 1.3369 0.03403 0.02447 -0.0614 0.0185 1.0000
10.000 1.3431 0.03570 0.02641 -0.0587 0.0177 1.0000
10.250 1.3464 0.03753 0.02849 -0.0559 0.0171 1.0000
10.500 1.3474 0.03954 0.03082 -0.0532 0.0165 1.0000
10.750 1.3466 0.04177 0.03329 -0.0507 0.0160 1.0000
11.000 1.3442 0.04425 0.03600 -0.0485 0.0154 1.0000
11.250 1.3405 0.04699 0.03897 -0.0467 0.0147 1.0000
11.500 1.3356 0.05003 0.04222 -0.0453 0.0142 1.0000
11.750 1.3293 0.05341 0.04581 -0.0444 0.0137 1.0000
12.000 1.3221 0.05707 0.04968 -0.0438 0.0133 1.0000
12.250 1.3143 0.06100 0.05382 -0.0435 0.0131 1.0000
12.500 1.3060 0.06519 0.05822 -0.0436 0.0129 1.0000
12.750 1.2974 0.06963 0.06286 -0.0441 0.0128 1.0000
13.000 1.2878 0.07438 0.06782 -0.0449 0.0127 1.0000
13.250 1.2775 0.07946 0.07310 -0.0461 0.0126 1.0000
13.500 1.2665 0.08489 0.07874 -0.0478 0.0125 1.0000
13.750 1.2548 0.09071 0.08476 -0.0500 0.0125 1.0000
14.000 1.2424 0.09698 0.09124 -0.0527 0.0125 1.0000
14.250 1.2297 0.10369 0.09814 -0.0560 0.0125 1.0000
14.500 1.2167 0.11081 0.10545 -0.0597 0.0125 1.0000
14.750 1.2034 0.11839 0.11320 -0.0638 0.0125 1.0000
15.000 1.1901 0.12631 0.12130 -0.0683 0.0126 1.0000
15.250 1.1765 0.13463 0.12978 -0.0731 0.0127 1.0000
15.500 1.1623 0.14351 0.13881 -0.0783 0.0129 1.0000
15.750 1.1477 0.15298 0.14842 -0.0839 0.0131 1.0000
16.000 1.1319 0.16349 0.15907 -0.0900 0.0134 1.0000
16.250 1.1132 0.17595 0.17162 -0.0969 0.0139 1.0000
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Polar data table (+)
Polar graphs
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