XFOIL Version 6.96 Calculated polar for: FX 66-S-196 V1 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -17.000 -0.6236 0.10159 0.09948 -0.0616 1.0000 0.0087 -16.750 -0.6450 0.09298 0.09074 -0.0666 1.0000 0.0087 -16.500 -0.6748 0.08351 0.08111 -0.0721 1.0000 0.0086 -16.250 -0.6950 0.07613 0.07357 -0.0765 1.0000 0.0086 -16.000 -0.7070 0.06938 0.06667 -0.0810 0.9626 0.0086 -15.750 -0.6933 0.06170 0.05870 -0.0929 0.9397 0.0087 -15.500 -0.6706 0.05461 0.05130 -0.1058 0.9196 0.0087 -15.250 -0.6601 0.04787 0.04415 -0.1164 0.8908 0.0090 -15.000 -0.6538 0.04435 0.04033 -0.1202 0.8589 0.0090 -14.750 -0.6527 0.04170 0.03749 -0.1215 0.8327 0.0093 -14.500 -0.6524 0.03932 0.03493 -0.1220 0.8150 0.0093 -14.250 -0.6504 0.03727 0.03271 -0.1222 0.7974 0.0093 -14.000 -0.6440 0.03566 0.03101 -0.1224 0.7821 0.0096 -13.750 -0.6384 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1.3522 0.08188 0.07646 -0.0668 0.0087 1.0000 16.000 1.3555 0.08444 0.07908 -0.0667 0.0084 1.0000 16.250 1.3612 0.08671 0.08140 -0.0666 0.0081 1.0000 16.500 1.3649 0.08925 0.08399 -0.0665 0.0077 1.0000 16.750 1.3676 0.09186 0.08665 -0.0664 0.0075 1.0000 17.000 1.3708 0.09440 0.08923 -0.0663 0.0072 1.0000 17.250 1.3724 0.09719 0.09208 -0.0663 0.0070 1.0000 17.500 1.3706 0.10038 0.09532 -0.0663 0.0067 1.0000 17.750 1.3673 0.10378 0.09879 -0.0663 0.0065 1.0000