XFOIL Version 6.96 Calculated polar for: FX S 03-182 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 ------ -------- --------- --------- -------- -------- -------- -12.750 -0.3040 0.08702 0.08470 -0.1025 0.9472 0.0208 -12.500 -0.3185 0.07516 0.07263 -0.1162 0.9385 0.0209 -12.250 -0.3293 0.06564 0.06287 -0.1278 0.9251 0.0208 -12.000 -0.5012 0.03420 0.02949 -0.1329 0.8945 0.0081 -11.750 -0.4907 0.03276 0.02781 -0.1327 0.8739 0.0079 -11.500 -0.4811 0.03169 0.02655 -0.1318 0.8575 0.0078 -11.250 -0.4763 0.02906 0.02362 -0.1301 0.8449 0.0076 -11.000 -0.4660 0.02771 0.02207 -0.1289 0.8336 0.0075 -10.750 -0.4544 0.02619 0.02032 -0.1274 0.8240 0.0073 -10.500 -0.4423 0.02493 0.01889 -0.1259 0.8146 0.0072 -10.250 -0.4299 0.02389 0.01769 -0.1245 0.8063 0.0070 -10.000 -0.4185 0.02287 0.01653 -0.1229 0.7984 0.0070 -9.750 -0.4060 0.02203 0.01556 -0.1215 0.7910 0.0069 -9.500 -0.3942 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0.02365 -0.0581 0.0126 0.7435 10.000 1.1045 0.03265 0.02481 -0.0573 0.0119 0.7473 10.250 1.1171 0.03380 0.02601 -0.0565 0.0111 0.7516 10.500 1.1301 0.03495 0.02722 -0.0558 0.0104 0.7561 10.750 1.1430 0.03609 0.02844 -0.0551 0.0099 0.7604 11.000 1.1559 0.03728 0.02970 -0.0544 0.0095 0.7657 11.250 1.1699 0.03841 0.03089 -0.0539 0.0092 0.7713 11.500 1.1834 0.03955 0.03213 -0.0534 0.0088 0.7769 11.750 1.1968 0.04076 0.03342 -0.0529 0.0082 0.7835 12.000 1.2096 0.04201 0.03477 -0.0524 0.0071 0.7898 12.250 1.2196 0.04356 0.03638 -0.0518 0.0041 0.7972 12.500 1.2271 0.04534 0.03820 -0.0510 0.0037 0.8044 12.750 1.2365 0.04700 0.03993 -0.0504 0.0035 0.8131 13.000 1.2461 0.04860 0.04164 -0.0498 0.0034 0.8223 13.250 1.2559 0.05024 0.04339 -0.0493 0.0033 0.8331 13.500 1.2650 0.05192 0.04519 -0.0488 0.0032 0.8459 13.750 1.2731 0.05366 0.04707 -0.0481 0.0032 0.8617 14.000 1.2795 0.05541 0.04898 -0.0472 0.0031 0.8853 14.250 1.2795 0.05684 0.05062 -0.0450 0.0030 1.0000 14.500 1.2876 0.05900 0.05286 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0.0027 1.0000