BOEING BACXXX AIRFOIL (bacxxx-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: BOEING BACXXX AIRFOIL (bacxxx-il) Reynolds number: 100,000 Max Cl/Cd: 44.04 at α=3.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-bacxxx-il-100000.txt Download as CSV file: xf-bacxxx-il-100000.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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.500 -0.3660 0.11916 0.11428 -0.0300 1.0000 0.0858
-11.250 -0.3748 0.11599 0.11115 -0.0317 1.0000 0.0887
-11.000 -0.4931 0.11988 0.11469 -0.0268 1.0000 0.0791
-10.750 -0.4852 0.11667 0.11148 -0.0262 1.0000 0.0822
-10.500 -0.4849 0.11306 0.10791 -0.0273 1.0000 0.0854
-10.250 -0.4930 0.10900 0.10392 -0.0305 1.0000 0.0886
-10.000 -0.5145 0.10421 0.09924 -0.0367 1.0000 0.0901
-9.750 -0.5382 0.09898 0.09408 -0.0422 1.0000 0.0903
-9.500 -0.4992 0.09695 0.09204 -0.0331 1.0000 0.0956
-9.250 -0.5030 0.09331 0.08845 -0.0339 1.0000 0.0992
-9.000 -0.5200 0.08856 0.08379 -0.0372 1.0000 0.1016
-8.750 -0.5453 0.08399 0.07929 -0.0397 1.0000 0.1022
-8.500 -0.5733 0.08105 0.07638 -0.0383 1.0000 0.1025
-8.250 -0.6037 0.07841 0.07366 -0.0367 1.0000 0.1035
-8.000 -0.6379 0.07714 0.07213 -0.0337 1.0000 0.1045
-7.750 -0.6121 0.07109 0.06640 -0.0330 1.0000 0.1095
-7.500 -0.6185 0.06839 0.06365 -0.0309 1.0000 0.1144
-7.000 -0.6299 0.06215 0.05726 -0.0267 1.0000 0.1261
-6.750 -0.6398 0.05927 0.05414 -0.0244 1.0000 0.1363
-5.500 -0.6164 0.04626 0.04101 -0.0138 1.0000 0.2363
-5.250 -0.6042 0.04454 0.03962 -0.0103 1.0000 0.2719
-5.000 -0.5975 0.04271 0.03772 -0.0077 1.0000 0.3074
-4.750 -0.5849 0.04084 0.03583 -0.0052 1.0000 0.3352
-4.500 -0.5152 0.03473 0.02651 -0.0090 1.0000 0.1197
-4.250 -0.4785 0.03158 0.02274 -0.0068 1.0000 0.0753
-4.000 -0.4514 0.02851 0.01944 -0.0060 1.0000 0.0667
-3.750 -0.4247 0.02773 0.01823 -0.0044 1.0000 0.0614
-3.500 -0.3977 0.02567 0.01607 -0.0038 1.0000 0.0596
-3.250 -0.3713 0.02419 0.01447 -0.0030 1.0000 0.0587
-3.000 -0.3458 0.02297 0.01319 -0.0021 1.0000 0.0588
-2.750 -0.3213 0.02201 0.01220 -0.0012 1.0000 0.0606
-2.500 -0.2982 0.02079 0.01111 -0.0004 1.0000 0.0657
-2.250 -0.2753 0.01997 0.01031 0.0004 1.0000 0.0723
-2.000 -0.2509 0.01910 0.00933 0.0009 1.0000 0.0831
-1.750 -0.2311 0.01577 0.00893 0.0031 1.0000 0.7007
-1.500 -0.2280 0.01592 0.00950 0.0109 1.0000 0.8200
-1.250 -0.2212 0.01611 0.00980 0.0171 1.0000 0.8798
-1.000 -0.1954 0.01669 0.01037 0.0202 1.0000 0.9499
-0.750 -0.1366 0.01710 0.01052 0.0136 1.0000 0.9752
-0.500 -0.0755 0.01738 0.01060 0.0057 1.0000 0.9918
-0.250 -0.0615 0.01698 0.01011 0.0055 1.0000 1.0000
0.000 -0.0590 0.01668 0.00970 0.0078 1.0000 1.0000
0.250 -0.0411 0.01680 0.00969 0.0080 1.0000 1.0000
0.500 0.0161 0.01739 0.01014 0.0011 0.9888 1.0000
0.750 0.0731 0.01789 0.01053 -0.0056 0.9770 1.0000
1.000 0.1271 0.01822 0.01080 -0.0114 0.9642 1.0000
1.250 0.1779 0.01838 0.01092 -0.0163 0.9499 1.0000
1.500 0.2292 0.01838 0.01092 -0.0211 0.9351 1.0000
1.750 0.2795 0.01819 0.01077 -0.0253 0.9187 1.0000
2.000 0.3314 0.01781 0.01044 -0.0295 0.9016 1.0000
2.250 0.3872 0.01719 0.00991 -0.0341 0.8845 1.0000
2.500 0.4457 0.01630 0.00915 -0.0388 0.8672 1.0000
2.750 0.4845 0.01552 0.00845 -0.0396 0.8409 1.0000
3.000 0.5224 0.01473 0.00774 -0.0400 0.8046 1.0000
3.250 0.5646 0.01398 0.00692 -0.0411 0.7416 1.0000
3.500 0.6082 0.01381 0.00628 -0.0426 0.6413 1.0000
3.750 0.6321 0.01437 0.00629 -0.0413 0.5657 1.0000
4.000 0.6527 0.01503 0.00660 -0.0397 0.5167 1.0000
4.250 0.6730 0.01566 0.00699 -0.0383 0.4801 1.0000
4.500 0.6939 0.01627 0.00741 -0.0370 0.4503 1.0000
4.750 0.7149 0.01688 0.00787 -0.0358 0.4241 1.0000
5.000 0.7365 0.01751 0.00837 -0.0347 0.4010 1.0000
5.250 0.7579 0.01811 0.00892 -0.0336 0.3789 1.0000
5.500 0.7803 0.01878 0.00951 -0.0327 0.3595 1.0000
5.750 0.8031 0.01950 0.01012 -0.0318 0.3411 1.0000
6.000 0.8237 0.02008 0.01078 -0.0306 0.3214 1.0000
6.250 0.8442 0.02064 0.01132 -0.0294 0.3025 1.0000
6.500 0.8656 0.02126 0.01190 -0.0283 0.2855 1.0000
6.750 0.8879 0.02198 0.01262 -0.0274 0.2703 1.0000
7.000 0.9098 0.02270 0.01338 -0.0265 0.2554 1.0000
7.250 0.9307 0.02338 0.01405 -0.0254 0.2400 1.0000
7.500 0.9503 0.02400 0.01472 -0.0241 0.2244 1.0000
7.750 0.9696 0.02467 0.01548 -0.0228 0.2093 1.0000
8.000 0.9891 0.02546 0.01635 -0.0215 0.1944 1.0000
8.250 1.0079 0.02633 0.01732 -0.0201 0.1794 1.0000
8.500 1.0263 0.02728 0.01833 -0.0187 0.1648 1.0000
8.750 1.0458 0.02842 0.01949 -0.0175 0.1521 1.0000
9.000 1.0669 0.02970 0.02075 -0.0166 0.1411 1.0000
9.250 1.0843 0.03113 0.02251 -0.0150 0.1315 1.0000
9.500 1.1031 0.03275 0.02427 -0.0137 0.1230 1.0000
9.750 1.1236 0.03415 0.02557 -0.0131 0.1139 1.0000
10.000 1.1329 0.03536 0.02724 -0.0103 0.1058 1.0000
10.250 1.1467 0.03681 0.02876 -0.0087 0.0974 1.0000
10.500 1.1592 0.03769 0.02966 -0.0069 0.0887 1.0000
10.750 1.1642 0.03945 0.03181 -0.0039 0.0814 1.0000
11.000 1.1749 0.04117 0.03352 -0.0021 0.0741 1.0000
11.250 1.1752 0.04283 0.03552 0.0013 0.0677 1.0000
11.500 1.1796 0.04534 0.03809 0.0036 0.0623 1.0000
11.750 1.1703 0.04819 0.04140 0.0077 0.0595 1.0000
12.000 1.1626 0.05089 0.04438 0.0112 0.0568 1.0000
12.250 1.1631 0.05312 0.04666 0.0133 0.0536 1.0000
12.500 1.1520 0.05761 0.05130 0.0155 0.0518 1.0000
12.750 1.1311 0.06090 0.05493 0.0181 0.0513 1.0000
13.000 1.1089 0.06494 0.05927 0.0195 0.0510 1.0000
13.250 1.0853 0.06976 0.06435 0.0196 0.0510 1.0000
13.500 1.0608 0.07542 0.07020 0.0182 0.0511 1.0000
13.750 1.0358 0.08199 0.07695 0.0155 0.0515 1.0000
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Polar data table (+)
Polar graphs
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