BOEING 103 AIRFOIL (boe103-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: BOEING 103 AIRFOIL (boe103-il) Reynolds number: 50,000 Max Cl/Cd: 35.02 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-boe103-il-50000-n5.txt Download as CSV file: xf-boe103-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: BOEING 103 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
------ -------- --------- --------- -------- -------- --------
-10.750 -0.3763 0.11732 0.10982 -0.0423 1.0000 0.0765
-10.500 -0.3720 0.11375 0.10629 -0.0424 1.0000 0.0762
-10.250 -0.3700 0.11021 0.10280 -0.0428 1.0000 0.0758
-10.000 -0.3689 0.10672 0.09937 -0.0430 1.0000 0.0753
-9.750 -0.3689 0.10328 0.09599 -0.0432 1.0000 0.0744
-9.500 -0.3712 0.09979 0.09257 -0.0434 1.0000 0.0733
-9.250 -0.3767 0.09620 0.08907 -0.0438 1.0000 0.0723
-9.000 -0.3857 0.09253 0.08551 -0.0441 1.0000 0.0713
-8.750 -0.3993 0.08883 0.08193 -0.0444 1.0000 0.0703
-8.500 -0.4200 0.08513 0.07837 -0.0443 1.0000 0.0694
-8.250 -0.4463 0.08061 0.07398 -0.0450 1.0000 0.0682
-7.750 -0.5129 0.06804 0.06112 -0.0467 1.0000 0.0653
-7.500 -0.5103 0.06631 0.05949 -0.0446 1.0000 0.0663
-7.250 -0.5122 0.06379 0.05691 -0.0431 1.0000 0.0671
-7.000 -0.5138 0.06085 0.05385 -0.0419 1.0000 0.0679
-6.750 -0.5136 0.05759 0.05040 -0.0409 1.0000 0.0685
-6.500 -0.5107 0.05412 0.04665 -0.0402 1.0000 0.0689
-6.250 -0.4860 0.04944 0.04145 -0.0436 0.9941 0.0692
-6.000 -0.4584 0.04555 0.03703 -0.0464 0.9874 0.0700
-5.750 -0.4285 0.04254 0.03351 -0.0488 0.9810 0.0723
-5.500 -0.3965 0.03965 0.03000 -0.0510 0.9748 0.0749
-5.250 -0.3652 0.03719 0.02695 -0.0525 0.9682 0.0763
-5.000 -0.3307 0.03530 0.02480 -0.0545 0.9625 0.0779
-4.750 -0.2997 0.03393 0.02327 -0.0556 0.9552 0.0801
-4.500 -0.2642 0.03278 0.02190 -0.0574 0.9491 0.0843
-4.250 -0.2328 0.03166 0.02047 -0.0581 0.9416 0.0888
-4.000 -0.1977 0.03064 0.01942 -0.0597 0.9354 0.0928
-3.750 -0.1663 0.02982 0.01851 -0.0604 0.9280 0.0980
-3.500 -0.1321 0.02903 0.01759 -0.0615 0.9211 0.1057
-3.250 -0.0977 0.02837 0.01691 -0.0629 0.9143 0.1200
-3.000 -0.0668 0.02769 0.01629 -0.0638 0.9063 0.1429
-2.750 -0.0278 0.02684 0.01556 -0.0661 0.9012 0.1878
-2.500 -0.0032 0.02624 0.01522 -0.0661 0.8914 0.2434
-2.250 0.0301 0.02499 0.01505 -0.0678 0.8860 0.4112
-2.000 0.0466 0.02422 0.01527 -0.0648 0.8766 0.6250
-1.750 0.0786 0.02385 0.01534 -0.0633 0.8711 0.8063
-1.500 0.1511 0.02378 0.01510 -0.0704 0.8686 0.9560
-1.250 0.1933 0.02391 0.01494 -0.0738 0.8587 1.0000
-1.000 0.2304 0.02386 0.01461 -0.0756 0.8522 1.0000
-0.750 0.2495 0.02408 0.01461 -0.0744 0.8401 1.0000
-0.500 0.2776 0.02419 0.01452 -0.0746 0.8309 1.0000
-0.250 0.3079 0.02425 0.01438 -0.0750 0.8218 1.0000
0.000 0.3320 0.02440 0.01437 -0.0744 0.8101 1.0000
0.250 0.3680 0.02425 0.01406 -0.0754 0.8014 1.0000
0.500 0.3940 0.02429 0.01397 -0.0748 0.7890 1.0000
0.750 0.4186 0.02438 0.01394 -0.0740 0.7764 1.0000
1.000 0.4481 0.02436 0.01380 -0.0739 0.7658 1.0000
1.250 0.4785 0.02432 0.01366 -0.0740 0.7557 1.0000
1.500 0.5012 0.02452 0.01378 -0.0729 0.7433 1.0000
1.750 0.5283 0.02459 0.01378 -0.0725 0.7323 1.0000
2.000 0.5599 0.02451 0.01363 -0.0726 0.7226 1.0000
2.250 0.5823 0.02473 0.01381 -0.0715 0.7096 1.0000
2.500 0.6073 0.02487 0.01391 -0.0707 0.6974 1.0000
2.750 0.6370 0.02484 0.01383 -0.0705 0.6865 1.0000
3.000 0.6646 0.02488 0.01384 -0.0700 0.6744 1.0000
3.250 0.6880 0.02507 0.01403 -0.0689 0.6606 1.0000
3.500 0.7127 0.02523 0.01418 -0.0680 0.6470 1.0000
3.750 0.7385 0.02536 0.01429 -0.0673 0.6335 1.0000
4.000 0.7651 0.02545 0.01437 -0.0666 0.6198 1.0000
4.250 0.7920 0.02554 0.01445 -0.0660 0.6057 1.0000
4.500 0.8181 0.02566 0.01455 -0.0652 0.5909 1.0000
4.750 0.8422 0.02586 0.01475 -0.0642 0.5751 1.0000
5.000 0.8654 0.02610 0.01500 -0.0631 0.5587 1.0000
5.250 0.8870 0.02640 0.01529 -0.0618 0.5414 1.0000
5.500 0.9074 0.02673 0.01563 -0.0603 0.5231 1.0000
5.750 0.9265 0.02705 0.01596 -0.0586 0.5031 1.0000
6.000 0.9457 0.02729 0.01615 -0.0568 0.4817 1.0000
6.250 0.9640 0.02756 0.01633 -0.0549 0.4598 1.0000
6.500 0.9794 0.02801 0.01675 -0.0528 0.4371 1.0000
6.750 0.9964 0.02845 0.01711 -0.0509 0.4161 1.0000
7.000 1.0130 0.02899 0.01758 -0.0491 0.3968 1.0000
7.250 1.0287 0.02964 0.01821 -0.0473 0.3786 1.0000
7.500 1.0442 0.03035 0.01890 -0.0455 0.3612 1.0000
7.750 1.0587 0.03112 0.01967 -0.0436 0.3439 1.0000
8.000 1.0714 0.03195 0.02048 -0.0416 0.3269 1.0000
8.250 1.0829 0.03286 0.02137 -0.0394 0.3103 1.0000
8.500 1.0939 0.03386 0.02236 -0.0374 0.2941 1.0000
8.750 1.1048 0.03493 0.02347 -0.0354 0.2788 1.0000
9.000 1.1156 0.03607 0.02462 -0.0335 0.2644 1.0000
9.250 1.1264 0.03727 0.02580 -0.0317 0.2511 1.0000
9.500 1.1371 0.03851 0.02703 -0.0300 0.2388 1.0000
9.750 1.1473 0.03988 0.02848 -0.0284 0.2269 1.0000
10.000 1.1581 0.04127 0.02993 -0.0269 0.2167 1.0000
10.250 1.1701 0.04254 0.03114 -0.0255 0.2083 1.0000
10.500 1.1807 0.04412 0.03287 -0.0242 0.1997 1.0000
10.750 1.1959 0.04534 0.03404 -0.0231 0.1935 1.0000
11.000 1.2080 0.04705 0.03595 -0.0221 0.1874 1.0000
11.250 1.2224 0.04853 0.03755 -0.0211 0.1822 1.0000
11.500 1.2404 0.04981 0.03881 -0.0203 0.1777 1.0000
11.750 1.2463 0.05199 0.04127 -0.0191 0.1731 1.0000
12.000 1.2547 0.05391 0.04335 -0.0181 0.1688 1.0000
12.250 1.2710 0.05528 0.04474 -0.0173 0.1649 1.0000
12.500 1.2776 0.05757 0.04722 -0.0163 0.1617 1.0000
12.750 1.2742 0.06062 0.05058 -0.0152 0.1588 1.0000
13.000 1.2714 0.06362 0.05381 -0.0144 0.1557 1.0000
13.250 1.2742 0.06608 0.05640 -0.0136 0.1526 1.0000
13.500 1.2951 0.06703 0.05727 -0.0130 0.1490 1.0000
13.750 1.2691 0.07230 0.06293 -0.0127 0.1471 1.0000
14.000 1.2382 0.07872 0.06967 -0.0134 0.1454 1.0000
14.250 1.1944 0.08759 0.07884 -0.0159 0.1440 1.0000
|
Polar data table (+)
Polar graphs
<< Back to BOEING 103 AIRFOIL (boe103-il)