BOEING 106 AIRFOIL (boe106-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: BOEING 106 AIRFOIL (boe106-il) Reynolds number: 50,000 Max Cl/Cd: 32.53 at α=8.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-boe106-il-50000-n5.txt Download as CSV file: xf-boe106-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: BOEING 106 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 -0.3919 0.09940 0.09213 -0.0456 1.0000 0.0749
-9.750 -0.3975 0.09507 0.08788 -0.0475 1.0000 0.0750
-9.500 -0.4059 0.09065 0.08355 -0.0493 1.0000 0.0751
-9.250 -0.4177 0.08624 0.07925 -0.0510 1.0000 0.0750
-9.000 -0.4355 0.08204 0.07516 -0.0521 1.0000 0.0746
-8.750 -0.4590 0.07864 0.07188 -0.0512 1.0000 0.0741
-8.500 -0.4814 0.07527 0.06856 -0.0497 1.0000 0.0738
-8.250 -0.5034 0.07202 0.06532 -0.0477 1.0000 0.0735
-8.000 -0.5233 0.06874 0.06200 -0.0456 1.0000 0.0733
-7.750 -0.5399 0.06534 0.05850 -0.0435 1.0000 0.0731
-7.500 -0.5525 0.06184 0.05483 -0.0414 1.0000 0.0730
-7.250 -0.5607 0.05830 0.05104 -0.0396 1.0000 0.0731
-7.000 -0.5554 0.05397 0.04629 -0.0401 0.9965 0.0738
-6.750 -0.5360 0.04906 0.04064 -0.0429 0.9885 0.0761
-6.500 -0.5109 0.04560 0.03671 -0.0450 0.9816 0.0780
-6.250 -0.4825 0.04308 0.03389 -0.0467 0.9746 0.0795
-6.000 -0.4540 0.04082 0.03130 -0.0482 0.9674 0.0816
-5.750 -0.4223 0.03875 0.02881 -0.0501 0.9607 0.0854
-5.500 -0.3934 0.03669 0.02617 -0.0510 0.9530 0.0885
-5.250 -0.3594 0.03489 0.02409 -0.0527 0.9467 0.0910
-5.000 -0.3288 0.03365 0.02273 -0.0537 0.9392 0.0943
-4.750 -0.2949 0.03252 0.02133 -0.0551 0.9322 0.1001
-4.500 -0.2606 0.03138 0.01998 -0.0564 0.9257 0.1054
-4.250 -0.2294 0.03046 0.01901 -0.0572 0.9178 0.1107
-4.000 -0.1895 0.02954 0.01793 -0.0594 0.9126 0.1205
-3.750 -0.1628 0.02888 0.01722 -0.0593 0.9033 0.1310
-3.500 -0.1237 0.02802 0.01631 -0.0615 0.8976 0.1480
-3.250 -0.0958 0.02728 0.01566 -0.0618 0.8887 0.1706
-3.000 -0.0619 0.02639 0.01498 -0.0632 0.8821 0.2116
-2.750 -0.0376 0.02565 0.01463 -0.0630 0.8727 0.2767
-2.500 -0.0101 0.02450 0.01431 -0.0631 0.8658 0.4281
-2.250 0.0055 0.02364 0.01459 -0.0594 0.8566 0.6624
-2.000 0.0542 0.02350 0.01474 -0.0607 0.8516 0.8449
-1.750 0.1283 0.02353 0.01447 -0.0680 0.8485 0.9348
-1.500 0.2281 0.02337 0.01393 -0.0817 0.8453 1.0000
-1.250 0.2503 0.02334 0.01370 -0.0810 0.8342 1.0000
-1.000 0.2684 0.02343 0.01361 -0.0796 0.8220 1.0000
-0.750 0.2918 0.02344 0.01343 -0.0790 0.8117 1.0000
-0.500 0.3167 0.02342 0.01325 -0.0784 0.8016 1.0000
-0.250 0.3348 0.02358 0.01326 -0.0768 0.7893 1.0000
0.000 0.3580 0.02363 0.01317 -0.0759 0.7789 1.0000
0.250 0.3837 0.02362 0.01301 -0.0753 0.7691 1.0000
0.500 0.4020 0.02383 0.01311 -0.0737 0.7569 1.0000
0.750 0.4252 0.02391 0.01308 -0.0727 0.7465 1.0000
1.000 0.4507 0.02393 0.01299 -0.0720 0.7368 1.0000
1.250 0.4692 0.02420 0.01317 -0.0703 0.7249 1.0000
1.500 0.4935 0.02429 0.01318 -0.0695 0.7152 1.0000
1.750 0.5176 0.02440 0.01321 -0.0686 0.7052 1.0000
2.000 0.5369 0.02471 0.01347 -0.0671 0.6937 1.0000
2.250 0.5645 0.02474 0.01342 -0.0667 0.6854 1.0000
2.500 0.5847 0.02504 0.01369 -0.0653 0.6742 1.0000
2.750 0.6055 0.02536 0.01399 -0.0641 0.6635 1.0000
3.000 0.6344 0.02537 0.01395 -0.0638 0.6556 1.0000
3.250 0.6524 0.02584 0.01442 -0.0623 0.6440 1.0000
3.500 0.6759 0.02611 0.01468 -0.0615 0.6344 1.0000
3.750 0.7012 0.02630 0.01487 -0.0608 0.6253 1.0000
4.000 0.7207 0.02679 0.01538 -0.0596 0.6147 1.0000
4.250 0.7510 0.02680 0.01537 -0.0595 0.6074 1.0000
4.500 0.7678 0.02744 0.01608 -0.0580 0.5961 1.0000
4.750 0.7912 0.02782 0.01648 -0.0572 0.5871 1.0000
5.000 0.8156 0.02815 0.01686 -0.0565 0.5782 1.0000
5.250 0.8350 0.02876 0.01753 -0.0554 0.5684 1.0000
5.500 0.8634 0.02894 0.01774 -0.0551 0.5609 1.0000
5.750 0.8796 0.02975 0.01866 -0.0537 0.5505 1.0000
6.000 0.9111 0.02980 0.01874 -0.0537 0.5437 1.0000
6.250 0.9247 0.03070 0.01976 -0.0520 0.5323 1.0000
6.500 0.9456 0.03111 0.02027 -0.0508 0.5213 1.0000
6.750 0.9717 0.03107 0.02025 -0.0499 0.5091 1.0000
7.000 0.9964 0.03101 0.02021 -0.0488 0.4956 1.0000
7.250 1.0123 0.03144 0.02072 -0.0468 0.4812 1.0000
7.500 1.0274 0.03203 0.02143 -0.0450 0.4680 1.0000
7.750 1.0440 0.03254 0.02204 -0.0433 0.4553 1.0000
8.000 1.0620 0.03294 0.02253 -0.0417 0.4425 1.0000
8.250 1.0797 0.03329 0.02297 -0.0400 0.4289 1.0000
8.500 1.0956 0.03368 0.02344 -0.0381 0.4146 1.0000
8.750 1.1081 0.03420 0.02404 -0.0359 0.3993 1.0000
9.000 1.1177 0.03483 0.02475 -0.0334 0.3833 1.0000
9.250 1.1236 0.03559 0.02557 -0.0305 0.3666 1.0000
9.500 1.1284 0.03651 0.02658 -0.0278 0.3491 1.0000
9.750 1.1324 0.03757 0.02770 -0.0252 0.3305 1.0000
10.000 1.1364 0.03874 0.02889 -0.0229 0.3113 1.0000
10.250 1.1403 0.04001 0.03013 -0.0207 0.2918 1.0000
10.500 1.1404 0.04177 0.03195 -0.0187 0.2718 1.0000
10.750 1.1413 0.04362 0.03381 -0.0169 0.2529 1.0000
11.000 1.1428 0.04553 0.03568 -0.0154 0.2357 1.0000
11.250 1.1442 0.04757 0.03770 -0.0140 0.2206 1.0000
11.500 1.1459 0.04972 0.03982 -0.0128 0.2075 1.0000
11.750 1.1483 0.05191 0.04197 -0.0117 0.1963 1.0000
12.000 1.1518 0.05402 0.04401 -0.0106 0.1865 1.0000
12.250 1.1555 0.05625 0.04624 -0.0097 0.1773 1.0000
12.500 1.1602 0.05850 0.04854 -0.0089 0.1693 1.0000
12.750 1.1681 0.06039 0.05044 -0.0079 0.1620 1.0000
13.000 1.1735 0.06276 0.05294 -0.0072 0.1555 1.0000
13.250 1.1800 0.06497 0.05523 -0.0065 0.1490 1.0000
13.500 1.1894 0.06693 0.05723 -0.0057 0.1433 1.0000
13.750 1.1891 0.07013 0.06070 -0.0055 0.1385 1.0000
14.000 1.1979 0.07221 0.06286 -0.0049 0.1335 1.0000
14.250 1.2046 0.07464 0.06539 -0.0044 0.1291 1.0000
14.500 1.1932 0.07927 0.07036 -0.0050 0.1257 1.0000
14.750 1.1865 0.08341 0.07471 -0.0055 0.1221 1.0000
15.000 1.2025 0.08456 0.07584 -0.0046 0.1176 1.0000
15.250 1.1877 0.08993 0.08145 -0.0058 0.1152 1.0000
15.500 1.1579 0.09787 0.08971 -0.0089 0.1137 1.0000
15.750 1.1185 0.10825 0.10036 -0.0136 0.1127 1.0000
16.000 1.0393 0.12858 0.12088 -0.0249 0.1131 1.0000
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Polar data table (+)
Polar graphs
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