XFOIL Version 6.96 Calculated polar for: WORTMANN FX 05-188 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 ------ -------- --------- --------- -------- -------- -------- -13.750 -0.2514 0.07289 0.06983 -0.1244 0.8579 0.0415 -13.500 -0.2907 0.06340 0.06022 -0.1291 0.8548 0.0412 -13.250 -0.3247 0.05592 0.05260 -0.1322 0.8507 0.0409 -13.000 -0.3637 0.04951 0.04597 -0.1344 0.8460 0.0412 -12.750 -0.3917 0.04492 0.04122 -0.1351 0.8408 0.0413 -12.500 -0.4185 0.04123 0.03736 -0.1347 0.8352 0.0415 -12.250 -0.6037 0.03618 0.03051 -0.1311 0.8389 0.0259 -12.000 -0.6006 0.03466 0.02879 -0.1293 0.8325 0.0258 -11.750 -0.5936 0.03390 0.02786 -0.1278 0.8270 0.0256 -11.500 -0.5586 0.03116 0.02503 -0.1291 0.8243 0.0251 -11.250 -0.5403 0.02948 0.02321 -0.1286 0.8203 0.0250 -11.000 -0.5256 0.02825 0.02181 -0.1276 0.8163 0.0251 -10.750 -0.5009 0.02692 0.02033 -0.1277 0.8131 0.0251 -10.500 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1.1950 0.01987 0.01376 -0.0754 0.2553 0.7725 9.250 1.1858 0.02191 0.01558 -0.0716 0.2132 0.7744 9.500 1.1795 0.02393 0.01739 -0.0683 0.1751 0.7762 9.750 1.1775 0.02580 0.01910 -0.0658 0.1442 0.7780 10.000 1.1768 0.02769 0.02083 -0.0636 0.1148 0.7796 10.250 1.1780 0.02952 0.02254 -0.0617 0.0934 0.7812 10.500 1.1836 0.03114 0.02410 -0.0603 0.0806 0.7828 10.750 1.1898 0.03277 0.02569 -0.0590 0.0697 0.7841 11.000 1.1956 0.03443 0.02732 -0.0577 0.0567 0.7853 11.250 1.2013 0.03614 0.02898 -0.0565 0.0465 0.7865 11.500 1.2084 0.03781 0.03064 -0.0555 0.0412 0.7878 11.750 1.2154 0.03952 0.03235 -0.0545 0.0379 0.7891 12.000 1.2239 0.04117 0.03406 -0.0537 0.0360 0.7907 12.250 1.2327 0.04284 0.03578 -0.0530 0.0346 0.7924 12.500 1.2402 0.04468 0.03766 -0.0523 0.0334 0.7939 12.750 1.2452 0.04681 0.03983 -0.0516 0.0322 0.7954 13.000 1.2502 0.04901 0.04209 -0.0510 0.0314 0.7969 13.250 1.2586 0.05093 0.04409 -0.0506 0.0308 0.7983 13.500 1.2663 0.05298 0.04621 -0.0503 0.0301 0.7997 13.750 1.2737 0.05503 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