XFOIL Version 6.96 Calculated polar for: FX 60-157 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 ------ -------- --------- --------- -------- -------- -------- -12.250 -0.5420 0.09204 0.08692 -0.0616 1.0000 0.0226 -12.000 -0.5675 0.08552 0.08036 -0.0641 1.0000 0.0224 -11.750 -0.5907 0.08045 0.07524 -0.0653 1.0000 0.0222 -11.500 -0.6153 0.07591 0.07064 -0.0657 1.0000 0.0220 -11.250 -0.6413 0.07188 0.06652 -0.0652 1.0000 0.0218 -11.000 -0.6694 0.06814 0.06267 -0.0638 1.0000 0.0217 -10.750 -0.6889 0.06352 0.05782 -0.0651 0.9969 0.0217 -10.500 -0.7069 0.05907 0.05304 -0.0663 0.9918 0.0219 -10.250 -0.7258 0.05551 0.04913 -0.0655 0.9859 0.0222 -10.000 -0.7351 0.05199 0.04521 -0.0650 0.9806 0.0224 -9.750 -0.7375 0.04870 0.04150 -0.0639 0.9748 0.0228 -9.500 -0.7247 0.04665 0.03930 -0.0641 0.9714 0.0233 -9.250 -0.7120 0.04449 0.03690 -0.0638 0.9678 0.0239 -9.000 -0.7015 0.04246 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0.8320 5.750 0.8362 0.02057 0.01229 -0.0560 0.5131 0.8348 6.000 0.8616 0.02086 0.01251 -0.0560 0.4889 0.8374 6.250 0.8855 0.02122 0.01281 -0.0558 0.4652 0.8401 6.500 0.9026 0.02157 0.01313 -0.0541 0.4423 0.8426 6.750 0.9204 0.02198 0.01348 -0.0527 0.4198 0.8454 7.000 0.9388 0.02245 0.01392 -0.0515 0.3974 0.8483 7.250 0.9579 0.02296 0.01441 -0.0507 0.3757 0.8510 7.500 0.9775 0.02354 0.01493 -0.0500 0.3561 0.8540 7.750 0.9947 0.02412 0.01547 -0.0488 0.3380 0.8569 8.000 1.0100 0.02468 0.01605 -0.0472 0.3202 0.8597 8.250 1.0259 0.02530 0.01668 -0.0459 0.3027 0.8626 8.500 1.0420 0.02600 0.01736 -0.0448 0.2859 0.8655 8.750 1.0581 0.02675 0.01812 -0.0438 0.2684 0.8684 9.000 1.0742 0.02756 0.01893 -0.0429 0.2505 0.8713 9.250 1.0855 0.02833 0.01970 -0.0411 0.2352 0.8744 9.500 1.0976 0.02918 0.02054 -0.0395 0.2217 0.8779 9.750 1.1111 0.03006 0.02146 -0.0383 0.2080 0.8816 10.000 1.1258 0.03097 0.02245 -0.0374 0.1942 0.8852 10.250 1.1382 0.03192 0.02344 -0.0361 0.1824 0.8885 10.500 1.1493 0.03290 0.02448 -0.0347 0.1719 0.8921 10.750 1.1598 0.03401 0.02563 -0.0334 0.1612 0.8959 11.000 1.1733 0.03506 0.02682 -0.0325 0.1497 0.8999 11.250 1.1834 0.03621 0.02806 -0.0313 0.1404 0.9042 11.500 1.1897 0.03753 0.02941 -0.0297 0.1323 0.9093 11.750 1.1996 0.03881 0.03081 -0.0285 0.1231 0.9149 12.000 1.2062 0.04018 0.03229 -0.0271 0.1154 0.9211 12.250 1.2127 0.04162 0.03388 -0.0258 0.1061 0.9282 12.500 1.2168 0.04328 0.03561 -0.0245 0.0975 0.9367 12.750 1.2186 0.04519 0.03758 -0.0233 0.0902 0.9484 13.000 1.2192 0.04704 0.03956 -0.0221 0.0829 0.9811 13.250 1.2261 0.04939 0.04206 -0.0224 0.0749 1.0000 13.500 1.2301 0.05211 0.04488 -0.0226 0.0673 1.0000 13.750 1.2314 0.05518 0.04802 -0.0230 0.0621 1.0000 14.000 1.2293 0.05873 0.05163 -0.0236 0.0581 1.0000 14.250 1.2288 0.06226 0.05531 -0.0243 0.0544 1.0000 14.500 1.2249 0.06628 0.05945 -0.0253 0.0516 1.0000 14.750 1.2183 0.07080 0.06403 -0.0266 0.0496 1.0000 15.000 1.2173 0.07477 0.06822 -0.0279 0.0471 1.0000 15.250 1.2140 0.07914 0.07276 -0.0294 0.0451 1.0000 15.500 1.2093 0.08382 0.07759 -0.0313 0.0435 1.0000 15.750 1.2036 0.08874 0.08264 -0.0334 0.0423 1.0000 16.000 1.1988 0.09365 0.08766 -0.0356 0.0413 1.0000 16.250 1.1957 0.09834 0.09252 -0.0378 0.0402 1.0000