XFOIL Version 6.94 Calculated polar for: nabad 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.008 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4203 0.06349 0.04716 0.0167 0.9999 0.4823 -2.750 -0.3917 0.06004 0.04360 0.0150 0.9999 0.4930 -2.500 -0.3607 0.05693 0.04018 0.0124 0.9999 0.5082 -2.250 -0.3296 0.05361 0.03700 0.0116 0.9999 0.5342 -2.000 -0.2973 0.05041 0.03397 0.0108 0.9999 0.5734 -1.750 -0.2634 0.04660 0.03109 0.0113 0.9999 0.6428 -1.500 -0.0881 0.04153 0.02384 -0.0186 0.9999 1.0001 -1.250 -0.0549 0.04157 0.02217 -0.0180 0.9999 1.0001 -1.000 -0.0280 0.04156 0.02123 -0.0169 0.9999 1.0001 -0.750 -0.0025 0.04158 0.02061 -0.0157 0.9999 1.0001 -0.500 0.0221 0.04163 0.02020 -0.0147 0.9999 1.0001 -0.250 0.0463 0.04170 0.01996 -0.0136 0.9999 1.0001 0.000 0.0701 0.04182 0.01986 -0.0127 0.9999 1.0001 0.250 0.0938 0.04195 0.01982 -0.0117 0.9999 1.0001 0.500 0.1171 0.04213 0.01989 -0.0107 0.9999 1.0001 0.750 0.1402 0.04234 0.02007 -0.0098 0.9999 1.0001 1.000 0.1630 0.04260 0.02037 -0.0090 0.9999 1.0001 1.250 0.1853 0.04292 0.02081 -0.0082 0.9999 1.0001 1.500 0.2069 0.04333 0.02142 -0.0075 0.9999 1.0001 1.750 0.2273 0.04388 0.02225 -0.0070 0.9999 1.0001 2.000 0.2451 0.04469 0.02343 -0.0066 0.9999 1.0001 2.250 0.2565 0.04611 0.02521 -0.0066 0.9999 1.0001 2.500 0.2541 0.04878 0.02803 -0.0068 0.9999 1.0001 2.750 0.2444 0.05227 0.03133 -0.0075 0.9999 1.0001 3.000 0.2410 0.05537 0.03418 -0.0085 0.9999 1.0001 3.250 0.2427 0.05808 0.03666 -0.0093 0.9999 1.0001 3.500 0.2473 0.06058 0.03895 -0.0102 0.9999 1.0001 3.750 0.2537 0.06295 0.04113 -0.0109 0.9999 1.0001 4.000 0.2615 0.06523 0.04324 -0.0116 0.9999 1.0001 4.250 0.2703 0.06747 0.04533 -0.0122 0.9999 1.0001 4.500 0.2797 0.06968 0.04741 -0.0129 0.9999 1.0001 4.750 0.2898 0.07187 0.04949 -0.0134 0.9999 1.0001 5.000 0.3002 0.07405 0.05158 -0.0140 0.9999 1.0001