FX 63-147 AIRFOIL (fx63147-il) Xfoil prediction polar at RE=50,000 Ncrit=9
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Airfoil: FX 63-147 AIRFOIL (fx63147-il) Reynolds number: 50,000 Max Cl/Cd: 18.56 at α=13° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-fx63147-il-50000.txt Download as CSV file: xf-fx63147-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: FX 63-147 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.250 -0.3562 0.13806 0.13110 -0.0348 1.0001 0.2049
-11.000 -0.3337 0.13284 0.12587 -0.0334 1.0001 0.2123
-10.750 -0.3572 0.13257 0.12575 -0.0348 1.0001 0.2204
-10.500 -0.3309 0.12738 0.12055 -0.0330 1.0001 0.2315
-10.250 -0.3291 0.12434 0.11759 -0.0326 1.0001 0.2404
-10.000 -0.3497 0.12372 0.11710 -0.0330 1.0001 0.2509
-9.750 -0.3320 0.11963 0.11304 -0.0314 1.0001 0.2649
-9.500 -0.3214 0.11619 0.10964 -0.0302 1.0001 0.2773
-9.250 -0.3176 0.11333 0.10686 -0.0291 1.0001 0.2901
-9.000 -0.3158 0.11067 0.10427 -0.0279 1.0001 0.3041
-8.750 -0.3147 0.10813 0.10182 -0.0265 1.0001 0.3188
-8.500 -0.3131 0.10560 0.09938 -0.0249 1.0001 0.3342
-6.250 -0.3926 0.08578 0.08070 -0.0061 1.0001 0.3850
-6.000 -0.4479 0.08438 0.07954 -0.0029 1.0001 0.3632
-5.750 -0.4950 0.08312 0.07850 0.0004 1.0001 0.3613
-5.500 -0.5759 0.05998 0.05352 -0.0303 1.0001 0.1297
-5.250 -0.5622 0.05573 0.04882 -0.0305 1.0001 0.1146
-5.000 -0.5438 0.05229 0.04453 -0.0311 1.0001 0.1022
-4.750 -0.5275 0.04926 0.04139 -0.0310 1.0001 0.0999
-4.500 -0.5084 0.04660 0.03839 -0.0312 1.0001 0.0971
-4.250 -0.4868 0.04420 0.03553 -0.0314 1.0001 0.0942
-4.000 -0.4638 0.04219 0.03300 -0.0314 1.0001 0.0919
-3.750 -0.4306 0.04059 0.03081 -0.0328 0.9970 0.0903
-3.500 -0.3817 0.03892 0.02880 -0.0363 0.9872 0.0908
-3.250 -0.3411 0.03767 0.02748 -0.0381 0.9775 0.0940
-3.000 -0.3039 0.03696 0.02665 -0.0388 0.9676 0.1001
-2.750 -0.2680 0.03622 0.02591 -0.0393 0.9584 0.1129
-2.500 -0.0385 0.03815 0.03060 -0.0454 0.9725 0.9999
-2.250 -0.0038 0.03791 0.03001 -0.0498 0.9585 0.9999
-2.000 0.0292 0.03777 0.02956 -0.0536 0.9453 0.9999
-1.750 0.0626 0.03771 0.02920 -0.0572 0.9331 0.9999
-1.500 0.0874 0.03772 0.02898 -0.0590 0.9201 0.9999
-1.250 0.1089 0.03782 0.02890 -0.0600 0.9073 0.9999
-1.000 0.1293 0.03801 0.02890 -0.0608 0.8950 0.9999
-0.750 0.1529 0.03821 0.02892 -0.0619 0.8837 0.9999
-0.500 0.1863 0.03835 0.02887 -0.0644 0.8737 0.9999
-0.250 0.1933 0.03880 0.02921 -0.0626 0.8614 0.9999
0.000 0.2026 0.03930 0.02961 -0.0611 0.8504 0.9999
0.250 0.2305 0.03962 0.02978 -0.0624 0.8412 0.9999
0.500 0.2372 0.04021 0.03029 -0.0604 0.8303 0.9999
0.750 0.2382 0.04094 0.03096 -0.0576 0.8202 0.9999
1.000 0.2751 0.04125 0.03113 -0.0600 0.8121 0.9999
1.250 0.2571 0.04226 0.03212 -0.0544 0.8014 0.9999
1.500 0.2659 0.04300 0.03279 -0.0527 0.7928 0.9999
1.750 0.2772 0.04367 0.03340 -0.0513 0.7837 0.9999
2.000 0.2703 0.04461 0.03430 -0.0474 0.7745 0.9999
2.250 0.2963 0.04520 0.03482 -0.0479 0.7666 0.9999
2.500 0.2757 0.04633 0.03594 -0.0424 0.7580 0.9999
2.750 0.3083 0.04694 0.03648 -0.0437 0.7504 0.9999
3.000 0.2803 0.04821 0.03775 -0.0376 0.7428 0.9999
3.250 0.3081 0.04900 0.03849 -0.0383 0.7354 0.9999
3.500 0.2928 0.05041 0.03990 -0.0345 0.7292 0.9999
3.750 0.3108 0.05152 0.04098 -0.0345 0.7219 0.9999
4.000 0.3178 0.05294 0.04240 -0.0337 0.7153 0.9999
4.250 0.3278 0.05442 0.04389 -0.0335 0.7089 0.9999
4.500 0.3516 0.05581 0.04527 -0.0345 0.7016 0.9999
4.750 0.3558 0.05752 0.04701 -0.0340 0.6950 0.9999
5.000 0.3857 0.05891 0.04841 -0.0355 0.6866 0.9999
5.250 0.3860 0.06100 0.05054 -0.0352 0.6822 0.9999
5.500 0.4016 0.06292 0.05251 -0.0362 0.6772 0.9999
5.750 0.4229 0.06478 0.05441 -0.0373 0.6704 0.9999
6.000 0.4271 0.06721 0.05689 -0.0379 0.6688 0.9999
6.250 0.4377 0.06988 0.05963 -0.0392 0.6703 0.9999
6.500 0.4542 0.07276 0.06259 -0.0412 0.6732 0.9999
6.750 0.3744 0.07841 0.06837 -0.0408 0.7690 0.9999
7.000 0.3883 0.08035 0.07037 -0.0415 0.7576 0.9999
7.250 0.4091 0.08275 0.07282 -0.0430 0.7457 0.9999
7.500 0.4409 0.08595 0.07611 -0.0459 0.7340 0.9999
7.750 0.4594 0.08807 0.07831 -0.0470 0.7202 0.9999
8.000 0.4635 0.08959 0.07990 -0.0465 0.7084 0.9999
8.250 0.4771 0.09193 0.08232 -0.0473 0.6970 0.9999
8.500 0.5014 0.09501 0.08550 -0.0493 0.6862 0.9999
8.750 0.5292 0.09818 0.08882 -0.0515 0.6734 0.9999
9.000 0.5328 0.09982 0.09054 -0.0512 0.6601 0.9999
9.250 0.5448 0.10214 0.09295 -0.0517 0.6460 0.9999
9.500 0.5568 0.10463 0.09554 -0.0523 0.6326 0.9999
9.750 0.5702 0.10736 0.09839 -0.0532 0.6201 0.9999
10.000 0.5917 0.11074 0.10191 -0.0548 0.6087 0.9999
10.250 0.6155 0.11416 0.10547 -0.0564 0.5947 0.9999
10.500 0.6270 0.11669 0.10814 -0.0570 0.5809 0.9999
10.750 0.6251 0.11877 0.11031 -0.0569 0.5688 0.9999
11.000 0.6356 0.12154 0.11319 -0.0575 0.5543 0.9999
11.250 0.6510 0.12445 0.11624 -0.0584 0.5377 0.9999
11.500 0.6600 0.12736 0.11927 -0.0591 0.5237 0.9999
13.000 1.1807 0.06360 0.05801 -0.0208 0.2591 0.9999
13.250 1.1159 0.07722 0.07178 -0.0246 0.2614 0.9999
13.500 1.1854 0.06585 0.05981 -0.0177 0.1990 0.9999
13.750 1.1786 0.06934 0.06321 -0.0173 0.1792 0.9999
14.000 1.1896 0.07114 0.06472 -0.0158 0.1566 0.9999
14.250 1.1960 0.07414 0.06770 -0.0150 0.1426 0.9999
14.500 1.2157 0.07640 0.06988 -0.0137 0.1299 0.9999
14.750 1.1969 0.08224 0.07613 -0.0150 0.1275 0.9999
15.000 1.1762 0.08858 0.08279 -0.0168 0.1257 0.9999
15.250 1.1483 0.09632 0.09082 -0.0199 0.1256 0.9999
15.500 1.1103 0.10617 0.10091 -0.0246 0.1274 0.9999
15.750 1.0695 0.11780 0.11265 -0.0309 0.1296 0.9999
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Polar data table (+)
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