BOEING BACXXX AIRFOIL (bacxxx-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: BOEING BACXXX AIRFOIL (bacxxx-il) Reynolds number: 100,000 Max Cl/Cd: 42.69 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-bacxxx-il-100000-n5.txt Download as CSV file: xf-bacxxx-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: BOEING BACXXX 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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.5088 0.09352 0.08852 -0.0400 1.0000 0.0507
-9.500 -0.5161 0.08876 0.08380 -0.0423 1.0000 0.0511
-9.250 -0.5295 0.08421 0.07929 -0.0444 1.0000 0.0512
-9.000 -0.5486 0.08051 0.07559 -0.0449 1.0000 0.0513
-8.000 -0.6057 0.06672 0.06152 -0.0360 1.0000 0.0337
-7.750 -0.6218 0.06201 0.05633 -0.0320 1.0000 0.0262
-7.500 -0.6228 0.05874 0.05302 -0.0298 1.0000 0.0256
-7.250 -0.6235 0.05566 0.04976 -0.0274 1.0000 0.0250
-7.000 -0.6222 0.05260 0.04651 -0.0249 1.0000 0.0244
-6.750 -0.6184 0.04953 0.04320 -0.0224 1.0000 0.0237
-6.500 -0.6121 0.04649 0.03988 -0.0200 1.0000 0.0231
-6.250 -0.6032 0.04349 0.03655 -0.0177 1.0000 0.0225
-6.000 -0.5917 0.04060 0.03331 -0.0154 1.0000 0.0220
-5.750 -0.5776 0.03827 0.03066 -0.0135 1.0000 0.0225
-5.500 -0.5616 0.03621 0.02828 -0.0118 1.0000 0.0235
-5.250 -0.5437 0.03419 0.02591 -0.0101 1.0000 0.0244
-5.000 -0.5241 0.03216 0.02354 -0.0086 1.0000 0.0246
-4.750 -0.5029 0.03019 0.02123 -0.0073 1.0000 0.0245
-4.500 -0.4804 0.02842 0.01919 -0.0062 1.0000 0.0245
-4.250 -0.4566 0.02683 0.01737 -0.0052 1.0000 0.0247
-4.000 -0.4324 0.02543 0.01579 -0.0043 1.0000 0.0249
-3.750 -0.4039 0.02414 0.01435 -0.0043 0.9988 0.0254
-3.500 -0.3730 0.02299 0.01310 -0.0048 0.9966 0.0261
-3.250 -0.3424 0.02197 0.01202 -0.0055 0.9945 0.0269
-3.000 -0.3140 0.02096 0.01107 -0.0061 0.9916 0.0298
-2.750 -0.2829 0.02039 0.01042 -0.0072 0.9885 0.0342
-2.500 -0.2506 0.01971 0.00959 -0.0084 0.9859 0.0385
-2.250 -0.2199 0.01917 0.00890 -0.0092 0.9818 0.0483
-2.000 -0.1926 0.01641 0.00817 -0.0108 0.9797 0.4720
-1.750 -0.1643 0.01600 0.00854 -0.0100 0.9745 0.6799
-1.500 -0.1345 0.01599 0.00879 -0.0092 0.9666 0.7612
-1.250 -0.0995 0.01594 0.00865 -0.0102 0.9570 0.7881
-1.000 -0.0608 0.01585 0.00844 -0.0122 0.9474 0.8023
-0.750 -0.0208 0.01574 0.00825 -0.0145 0.9389 0.8172
-0.500 0.0101 0.01563 0.00809 -0.0148 0.9291 0.8336
-0.250 0.0486 0.01551 0.00794 -0.0167 0.9217 0.8501
0.000 0.0828 0.01539 0.00780 -0.0176 0.9119 0.8683
0.250 0.1183 0.01527 0.00767 -0.0188 0.9025 0.8875
0.500 0.1622 0.01511 0.00749 -0.0217 0.8953 0.9055
0.750 0.2000 0.01498 0.00736 -0.0235 0.8843 0.9241
1.000 0.2416 0.01481 0.00720 -0.0261 0.8729 0.9412
1.250 0.2841 0.01461 0.00701 -0.0288 0.8596 0.9569
1.500 0.3269 0.01439 0.00679 -0.0316 0.8427 0.9715
1.750 0.3710 0.01413 0.00654 -0.0347 0.8210 0.9840
2.000 0.4155 0.01388 0.00630 -0.0378 0.7903 0.9946
2.250 0.4597 0.01360 0.00593 -0.0406 0.7445 1.0000
2.500 0.4996 0.01340 0.00548 -0.0424 0.6811 1.0000
3.000 0.5493 0.01393 0.00528 -0.0407 0.5437 1.0000
3.250 0.5682 0.01435 0.00541 -0.0391 0.4952 1.0000
3.500 0.5870 0.01478 0.00559 -0.0375 0.4558 1.0000
3.750 0.6062 0.01517 0.00581 -0.0359 0.4226 1.0000
4.000 0.6258 0.01557 0.00608 -0.0345 0.3950 1.0000
4.250 0.6461 0.01595 0.00635 -0.0332 0.3734 1.0000
4.500 0.6671 0.01634 0.00666 -0.0321 0.3554 1.0000
4.750 0.6886 0.01671 0.00699 -0.0310 0.3379 1.0000
5.000 0.7101 0.01710 0.00738 -0.0300 0.3210 1.0000
5.250 0.7317 0.01749 0.00776 -0.0290 0.3049 1.0000
5.500 0.7534 0.01790 0.00816 -0.0280 0.2901 1.0000
5.750 0.7753 0.01833 0.00859 -0.0271 0.2766 1.0000
6.000 0.7970 0.01877 0.00907 -0.0261 0.2625 1.0000
6.250 0.8183 0.01923 0.00953 -0.0251 0.2473 1.0000
6.500 0.8394 0.01968 0.01000 -0.0241 0.2314 1.0000
6.750 0.8606 0.02016 0.01050 -0.0231 0.2171 1.0000
7.250 0.9028 0.02120 0.01167 -0.0212 0.1932 1.0000
7.500 0.9231 0.02175 0.01225 -0.0201 0.1800 1.0000
7.750 0.9430 0.02233 0.01286 -0.0190 0.1667 1.0000
8.000 0.9626 0.02295 0.01354 -0.0179 0.1547 1.0000
8.250 0.9826 0.02358 0.01432 -0.0168 0.1430 1.0000
8.500 1.0018 0.02428 0.01511 -0.0157 0.1315 1.0000
8.750 1.0200 0.02503 0.01595 -0.0144 0.1206 1.0000
9.000 1.0368 0.02589 0.01685 -0.0130 0.1111 1.0000
9.250 1.0533 0.02683 0.01788 -0.0115 0.1027 1.0000
9.500 1.0695 0.02787 0.01906 -0.0100 0.0953 1.0000
9.750 1.0841 0.02893 0.02025 -0.0083 0.0888 1.0000
10.000 1.0991 0.03004 0.02153 -0.0067 0.0822 1.0000
10.250 1.1108 0.03103 0.02257 -0.0048 0.0758 1.0000
10.500 1.1229 0.03203 0.02378 -0.0029 0.0687 1.0000
10.750 1.1300 0.03320 0.02496 -0.0006 0.0630 1.0000
11.000 1.1392 0.03444 0.02643 0.0015 0.0569 1.0000
11.250 1.1424 0.03600 0.02803 0.0039 0.0525 1.0000
11.500 1.1482 0.03764 0.02993 0.0060 0.0480 1.0000
11.750 1.1509 0.03935 0.03183 0.0082 0.0437 1.0000
12.000 1.1489 0.04146 0.03404 0.0103 0.0406 1.0000
12.250 1.1492 0.04366 0.03653 0.0120 0.0370 1.0000
12.500 1.1461 0.04611 0.03914 0.0135 0.0345 1.0000
12.750 1.1397 0.04899 0.04214 0.0145 0.0328 1.0000
13.000 1.1312 0.05241 0.04568 0.0152 0.0314 1.0000
13.250 1.1232 0.05618 0.04974 0.0152 0.0299 1.0000
13.500 1.1130 0.06047 0.05427 0.0145 0.0287 1.0000
13.750 1.1012 0.06533 0.05934 0.0128 0.0277 1.0000
14.000 1.0880 0.07082 0.06504 0.0103 0.0270 1.0000
14.250 1.0732 0.07700 0.07140 0.0071 0.0264 1.0000
14.500 1.0568 0.08380 0.07838 0.0033 0.0261 1.0000
14.750 1.0384 0.09128 0.08604 -0.0008 0.0260 1.0000
15.000 1.0168 0.09965 0.09459 -0.0054 0.0262 1.0000
15.250 0.9892 0.10971 0.10484 -0.0109 0.0268 1.0000
15.500 0.9535 0.12217 0.11746 -0.0176 0.0281 1.0000
15.750 0.9144 0.13645 0.13179 -0.0252 0.0294 1.0000
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Polar data table (+)
Polar graphs
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