XFOIL Version 6.94 Calculated polar for: DU 86-137/25 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.0081 0.01394 0.00409 -0.0001 0.3438 0.0786 0.500 0.0619 0.01384 0.00403 -0.0001 0.3435 0.0779 1.000 0.1160 0.01377 0.00400 -0.0002 0.3425 0.0764 1.500 0.1702 0.01366 0.00399 -0.0003 0.3398 0.0784 2.000 0.2226 0.01329 0.00397 -0.0003 0.3383 0.1717 2.500 0.2606 0.01130 0.00388 0.0018 0.3369 0.6729 3.000 0.3141 0.01131 0.00391 0.0021 0.3091 0.7260 3.500 0.3611 0.01206 0.00451 0.0034 0.3039 0.7612 4.000 0.4115 0.01479 0.00659 0.0041 0.2935 0.7911 4.500 0.4600 0.01733 0.00928 0.0060 0.2800 0.8268 5.000 0.4956 0.02001 0.01229 0.0111 0.2602 0.8607 6.000 0.5843 0.01946 0.01246 0.0147 0.2212 0.8775 6.500 0.6253 0.01668 0.00999 0.0176 0.1423 0.8843 7.000 0.6664 0.01733 0.01034 0.0195 0.0782 0.8903 7.500 0.7048 0.01848 0.01140 0.0215 0.0533 0.8971 8.000 0.7356 0.01980 0.01280 0.0248 0.0426 0.9039 8.500 0.7696 0.02098 0.01413 0.0275 0.0364 0.9115 9.000 0.7871 0.02297 0.01624 0.0324 0.0328 0.9209 9.500 0.8165 0.02462 0.01811 0.0358 0.0304 0.9309 10.000 0.8473 0.02635 0.01994 0.0382 0.0274 0.9424 10.500 0.8893 0.03013 0.02403 0.0387 0.0258 0.9544 11.000 0.9358 0.03413 0.02857 0.0374 0.0245 0.9725 11.500 0.9592 0.03925 0.03426 0.0380 0.0238 1.0000 12.000 0.9619 0.04413 0.03957 0.0399 0.0229 1.0000 12.500 0.9546 0.04953 0.04533 0.0412 0.0223 1.0000 13.000 0.9170 0.05832 0.05463 0.0415 0.0223 1.0000 13.500 0.8417 0.07377 0.07067 0.0358 0.0235 1.0000 14.000 0.7597 0.09904 0.09629 0.0183 0.0249 1.0000