BOEING HSNLF AIRFOIL (bacnlf-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: BOEING HSNLF AIRFOIL (bacnlf-il) Reynolds number: 500,000 Max Cl/Cd: 78.38 at α=2.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-bacnlf-il-500000.txt Download as CSV file: xf-bacnlf-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: BOEING HSNLF AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-17.250 -0.4327 0.17446 0.17220 -0.0170 1.0014 0.0083
-17.000 -0.4273 0.17033 0.16807 -0.0171 1.0014 0.0087
-5.250 -0.4503 0.02070 0.01495 -0.0461 0.9775 0.0119
-5.000 -0.4210 0.01729 0.01107 -0.0454 0.9769 0.0087
-4.750 -0.3894 0.01651 0.01020 -0.0463 0.9758 0.0084
-4.500 -0.3591 0.01511 0.00871 -0.0471 0.9749 0.0082
-4.250 -0.3276 0.01396 0.00746 -0.0483 0.9740 0.0082
-4.000 -0.2953 0.01295 0.00632 -0.0497 0.9732 0.0086
-3.750 -0.2613 0.01227 0.00553 -0.0514 0.9725 0.0104
-3.500 -0.2399 0.01171 0.00496 -0.0502 0.9682 0.0283
-3.250 -0.2102 0.01061 0.00454 -0.0518 0.9666 0.1925
-3.000 -0.1787 0.00901 0.00413 -0.0543 0.9662 0.5166
-2.750 -0.1470 0.00885 0.00416 -0.0554 0.9645 0.5912
-2.500 -0.1142 0.00887 0.00424 -0.0566 0.9630 0.6341
-2.250 -0.0811 0.00908 0.00453 -0.0575 0.9615 0.6770
-2.000 -0.0464 0.00919 0.00468 -0.0589 0.9603 0.6952
-1.750 -0.0201 0.00921 0.00467 -0.0587 0.9561 0.7022
-1.500 0.0112 0.00916 0.00458 -0.0596 0.9532 0.7060
-1.250 0.0444 0.00911 0.00453 -0.0610 0.9515 0.7094
-1.000 0.0785 0.00907 0.00448 -0.0626 0.9501 0.7132
-0.750 0.1138 0.00903 0.00442 -0.0644 0.9491 0.7171
-0.500 0.1489 0.00896 0.00436 -0.0662 0.9481 0.7203
-0.250 0.1835 0.00889 0.00432 -0.0678 0.9471 0.7236
0.000 0.2060 0.00893 0.00438 -0.0668 0.9415 0.7275
0.250 0.2387 0.00883 0.00430 -0.0680 0.9387 0.7314
0.500 0.2739 0.00863 0.00414 -0.0695 0.9364 0.7345
0.750 0.3104 0.00838 0.00395 -0.0712 0.9342 0.7381
1.000 0.3358 0.00827 0.00388 -0.0706 0.9270 0.7422
1.250 0.3715 0.00782 0.00345 -0.0717 0.9206 0.7459
1.500 0.3983 0.00729 0.00296 -0.0706 0.9049 0.7494
1.750 0.4249 0.00707 0.00278 -0.0701 0.8934 0.7535
2.000 0.4512 0.00692 0.00271 -0.0695 0.8807 0.7580
2.250 0.4772 0.00677 0.00259 -0.0688 0.8604 0.7620
2.500 0.5040 0.00666 0.00251 -0.0683 0.8350 0.7661
2.750 0.5283 0.00674 0.00239 -0.0671 0.7645 0.7708
3.000 0.5356 0.00790 0.00255 -0.0626 0.5481 0.7758
3.250 0.5460 0.00921 0.00301 -0.0595 0.3461 0.7808
3.500 0.5630 0.01025 0.00342 -0.0579 0.1934 0.7865
3.750 0.5833 0.01100 0.00379 -0.0569 0.1028 0.7917
4.000 0.6066 0.01147 0.00412 -0.0562 0.0634 0.7979
4.250 0.6306 0.01186 0.00447 -0.0557 0.0404 0.8044
4.500 0.6529 0.01251 0.00508 -0.0545 0.0169 0.8115
4.750 0.6765 0.01301 0.00565 -0.0537 0.0139 0.8190
5.000 0.6977 0.01381 0.00659 -0.0524 0.0119 0.8279
5.250 0.7205 0.01432 0.00723 -0.0514 0.0111 0.8378
5.500 0.7419 0.01505 0.00812 -0.0501 0.0105 0.8499
5.750 0.7628 0.01587 0.00910 -0.0487 0.0100 0.8652
6.250 0.8008 0.01774 0.01138 -0.0448 0.0091 0.9196
6.500 0.8258 0.01939 0.01330 -0.0441 0.0090 0.9986
6.750 0.8502 0.02268 0.01699 -0.0429 0.0096 0.9986
7.000 0.8671 0.02792 0.02284 -0.0403 0.0110 0.9986
7.750 0.8694 0.04962 0.04601 -0.0282 0.0188 0.9986
8.000 0.8724 0.05316 0.04982 -0.0254 0.0187 0.9986
8.250 0.8730 0.05667 0.05359 -0.0225 0.0187 0.9986
8.500 0.8713 0.06011 0.05726 -0.0197 0.0186 0.9986
8.750 0.8691 0.06323 0.06060 -0.0169 0.0184 0.9986
9.000 0.8753 0.06593 0.06356 -0.0147 0.0173 0.9986
9.250 0.8626 0.06983 0.06761 -0.0119 0.0169 0.9986
9.500 0.8453 0.07388 0.07176 -0.0099 0.0168 0.9986
9.750 0.8262 0.07854 0.07650 -0.0096 0.0168 0.9986
10.000 0.8067 0.08426 0.08231 -0.0116 0.0170 0.9986
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