XFOIL Version 6.96 Calculated polar for: FX60-100 10.0% smoothed 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -9.500 -0.3759 0.10095 0.09737 -0.0411 1.0000 0.0465 -9.250 -0.3789 0.09754 0.09402 -0.0432 1.0000 0.0467 -9.000 -0.3856 0.09235 0.08893 -0.0442 1.0000 0.0473 -8.750 -0.3720 0.09022 0.08678 -0.0403 1.0000 0.0488 -8.500 -0.3676 0.08809 0.08468 -0.0390 1.0000 0.0500 -8.250 -0.3682 0.08577 0.08241 -0.0383 1.0000 0.0513 -8.000 -0.3739 0.08345 0.08016 -0.0373 1.0000 0.0526 -7.750 -0.3856 0.08139 0.07818 -0.0356 1.0000 0.0537 -7.500 -0.4039 0.07973 0.07662 -0.0328 1.0000 0.0544 -7.250 -0.4237 0.07810 0.07508 -0.0300 1.0000 0.0548 -7.000 -0.4388 0.07516 0.07222 -0.0304 1.0000 0.0559 -6.500 -0.4414 0.05553 0.05223 -0.0597 0.9964 0.0605 -6.250 -0.4158 0.05411 0.05097 -0.0590 0.9937 0.0630 -6.000 -0.3765 0.04564 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0.0204 1.0000 13.500 1.4348 0.06223 0.05667 -0.0597 0.0198 1.0000 13.750 1.4276 0.06642 0.06113 -0.0599 0.0195 1.0000 14.000 1.4182 0.07115 0.06612 -0.0605 0.0192 1.0000 14.250 1.4064 0.07644 0.07168 -0.0617 0.0190 1.0000 14.500 1.3926 0.08228 0.07777 -0.0636 0.0189 1.0000 14.750 1.3769 0.08876 0.08452 -0.0662 0.0188 1.0000 15.000 1.3595 0.09593 0.09192 -0.0696 0.0188 1.0000 15.250 1.3405 0.10381 0.10005 -0.0738 0.0188 1.0000 15.500 1.3205 0.11244 0.10889 -0.0789 0.0189 1.0000 15.750 1.2994 0.12185 0.11852 -0.0849 0.0190 1.0000 16.000 1.2769 0.13216 0.12902 -0.0917 0.0192 1.0000 16.250 1.2536 0.14337 0.14040 -0.0994 0.0195 1.0000 16.500 1.2272 0.15618 0.15334 -0.1081 0.0198 1.0000