XFOIL Version 6.96 Calculated polar for: B-29 ROOT 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 ------ -------- --------- --------- -------- -------- -------- -19.750 -1.0866 0.07735 0.07245 -0.0655 1.0000 0.0543 -19.500 -1.0845 0.07457 0.06962 -0.0666 1.0000 0.0549 -19.250 -1.0824 0.07183 0.06682 -0.0676 1.0000 0.0555 -19.000 -1.0811 0.06901 0.06392 -0.0687 1.0000 0.0562 -18.750 -1.0801 0.06620 0.06102 -0.0696 1.0000 0.0568 -18.500 -1.0779 0.06358 0.05830 -0.0705 1.0000 0.0573 -18.250 -1.0772 0.06075 0.05535 -0.0714 1.0000 0.0579 -18.000 -1.0705 0.05890 0.05347 -0.0715 1.0000 0.0587 -17.750 -1.0628 0.05715 0.05174 -0.0716 1.0000 0.0595 -17.500 -1.0565 0.05515 0.04974 -0.0721 1.0000 0.0603 -17.250 -1.0505 0.05313 0.04768 -0.0725 1.0000 0.0610 -17.000 -1.0446 0.05110 0.04561 -0.0729 1.0000 0.0619 -16.750 -1.0384 0.04916 0.04360 -0.0731 1.0000 0.0627 -16.500 -1.0323 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0.5031 0.8037 2.750 0.3269 0.00979 0.00416 0.0081 0.4994 0.8180 3.000 0.3538 0.00984 0.00426 0.0084 0.4954 0.8314 3.250 0.3818 0.00992 0.00437 0.0084 0.4917 0.8438 3.500 0.4065 0.01005 0.00447 0.0090 0.4879 0.8573 3.750 0.4401 0.01020 0.00467 0.0078 0.4841 0.8675 4.000 0.4659 0.01029 0.00481 0.0082 0.4806 0.8790 4.250 0.4990 0.01044 0.00500 0.0071 0.4768 0.8879 4.500 0.5265 0.01059 0.00514 0.0072 0.4733 0.8977 4.750 0.5646 0.01085 0.00537 0.0050 0.4692 0.9035 5.000 0.5848 0.01101 0.00555 0.0065 0.4659 0.9141 5.250 0.6252 0.01124 0.00582 0.0038 0.4623 0.9174 5.500 0.6638 0.01147 0.00607 0.0015 0.4586 0.9210 5.750 0.6952 0.01169 0.00628 0.0007 0.4551 0.9266 6.000 0.7227 0.01193 0.00648 0.0007 0.4514 0.9328 6.250 0.7644 0.01223 0.00679 -0.0024 0.4476 0.9342 6.500 0.8026 0.01248 0.00709 -0.0047 0.4441 0.9361 6.750 0.8391 0.01274 0.00738 -0.0066 0.4405 0.9384 7.000 0.8714 0.01300 0.00764 -0.0077 0.4369 0.9424 7.250 0.8930 0.01325 0.00783 -0.0066 0.4325 0.9493 7.500 0.9307 0.01351 0.00815 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12.500 1.2864 0.02212 0.01713 0.0069 0.3016 1.0000 12.750 1.2787 0.02350 0.01849 0.0105 0.2928 1.0000 13.000 1.2754 0.02482 0.01981 0.0135 0.2823 1.0000 13.250 1.2684 0.02643 0.02141 0.0165 0.2714 1.0000 13.500 1.2581 0.02838 0.02332 0.0196 0.2600 1.0000 13.750 1.2479 0.03049 0.02540 0.0223 0.2467 1.0000 14.000 1.2376 0.03277 0.02764 0.0247 0.2325 1.0000 14.250 1.2220 0.03560 0.03041 0.0272 0.2164 1.0000 14.500 1.2071 0.03855 0.03330 0.0293 0.2015 1.0000 14.750 1.1877 0.04205 0.03673 0.0312 0.1856 1.0000 15.000 1.1761 0.04504 0.03968 0.0326 0.1738 1.0000 15.250 1.1639 0.04821 0.04281 0.0339 0.1621 1.0000 15.500 1.1538 0.05130 0.04589 0.0349 0.1523 1.0000 15.750 1.1432 0.05454 0.04910 0.0358 0.1437 1.0000 16.000 1.1349 0.05764 0.05218 0.0364 0.1345 1.0000 16.250 1.1300 0.06048 0.05501 0.0370 0.1266 1.0000 16.500 1.1220 0.06371 0.05823 0.0374 0.1198 1.0000 16.750 1.1185 0.06653 0.06104 0.0377 0.1133 1.0000 17.000 1.1135 0.06955 0.06406 0.0379 0.1080 1.0000 17.250 1.1104 0.07243 0.06694 0.0380 0.1030 1.0000 17.500 1.1061 0.07548 0.07000 0.0380 0.0989 1.0000 17.750 1.1065 0.07806 0.07260 0.0380 0.0951 1.0000 18.000 1.1024 0.08117 0.07571 0.0379 0.0918 1.0000 18.250 1.1033 0.08372 0.07828 0.0378 0.0888 1.0000 18.500 1.1032 0.08646 0.08106 0.0375 0.0860 1.0000 18.750 1.1021 0.08933 0.08393 0.0373 0.0835 1.0000