XFOIL Version 6.96 Calculated polar for: AH21 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.015 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -1.000 -0.0593 0.02821 0.00904 -0.0242 1.0000 1.0000 -0.900 -0.0510 0.02822 0.00890 -0.0238 1.0000 1.0000 -0.800 -0.0427 0.02823 0.00877 -0.0235 1.0000 1.0000 -0.700 -0.0344 0.02825 0.00865 -0.0231 1.0000 1.0000 -0.600 -0.0260 0.02828 0.00854 -0.0228 1.0000 1.0000 -0.500 -0.0176 0.02830 0.00845 -0.0225 1.0000 1.0000 -0.400 -0.0092 0.02834 0.00837 -0.0221 1.0000 1.0000 -0.300 -0.0008 0.02837 0.00830 -0.0218 1.0000 1.0000 -0.200 0.0076 0.02842 0.00824 -0.0215 1.0000 1.0000 -0.100 0.0160 0.02846 0.00819 -0.0212 1.0000 1.0000 0.000 0.0244 0.02851 0.00811 -0.0209 1.0000 1.0000 0.100 0.0329 0.02857 0.00808 -0.0206 1.0000 1.0000 0.200 0.0413 0.02863 0.00806 -0.0203 1.0000 1.0000 0.300 0.0497 0.02869 0.00805 -0.0200 1.0000 1.0000 0.400 0.0582 0.02876 0.00805 -0.0197 1.0000 1.0000 0.500 0.0666 0.02883 0.00807 -0.0194 1.0000 1.0000 0.600 0.0750 0.02891 0.00809 -0.0191 1.0000 1.0000 0.700 0.0834 0.02899 0.00813 -0.0188 1.0000 1.0000 0.800 0.0918 0.02908 0.00817 -0.0186 1.0000 1.0000 0.900 0.1002 0.02917 0.00823 -0.0183 1.0000 1.0000 1.000 0.1085 0.02927 0.00829 -0.0180 1.0000 1.0000 2.000 0.1911 0.03050 0.00950 -0.0154 1.0000 1.0000 3.000 0.2699 0.03235 0.01186 -0.0132 1.0000 1.0000 4.000 0.3422 0.03510 0.01565 -0.0115 1.0000 1.0000 5.000 0.4017 0.03949 0.02142 -0.0111 1.0000 1.0000 6.000 0.4324 0.04781 0.03130 -0.0149 1.0000 1.0000 8.000 0.7657 0.09203 0.07945 -0.0519 0.4417 1.0000 9.000 0.7359 0.11165 0.09871 -0.0597 0.4694 1.0000 10.000 0.5519 0.11709 0.10298 -0.0633 0.9398 1.0000