XFOIL Version 6.96 Calculated polar for: FX 61-163 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -11.000 -0.5533 0.06026 0.05608 -0.0681 1.0001 0.0102 -10.750 -0.5702 0.05639 0.05209 -0.0684 1.0001 0.0102 -10.500 -0.5861 0.05307 0.04865 -0.0678 1.0001 0.0102 -10.250 -0.6025 0.05020 0.04565 -0.0662 1.0001 0.0102 -10.000 -0.6128 0.04693 0.04217 -0.0659 0.9920 0.0102 -9.750 -0.6101 0.04292 0.03781 -0.0683 0.9597 0.0102 -9.500 -0.5898 0.03842 0.03286 -0.0731 0.9292 0.0103 -9.250 -0.5566 0.03462 0.02865 -0.0785 0.9029 0.0105 -9.000 -0.5194 0.03210 0.02579 -0.0832 0.8726 0.0109 -8.750 -0.4875 0.03053 0.02394 -0.0860 0.8418 0.0116 -8.500 -0.4622 0.02896 0.02202 -0.0869 0.8157 0.0123 -8.250 -0.4390 0.02738 0.02009 -0.0869 0.7953 0.0128 -8.000 -0.4162 0.02590 0.01829 -0.0863 0.7785 0.0130 -7.750 -0.3933 0.02463 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