XFOIL Version 6.94 Calculated polar for: GOA 2 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.018 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1746 0.03088 0.01678 -0.0313 1.0000 0.3424 -2.750 -0.1587 0.02641 0.01466 -0.0268 1.0000 0.5395 -2.500 -0.1123 0.02437 0.01160 -0.0258 1.0000 1.0000 -2.250 -0.0871 0.02434 0.01079 -0.0246 1.0000 1.0000 -2.000 -0.0626 0.02433 0.01014 -0.0235 1.0000 1.0000 -1.750 -0.0384 0.02434 0.00965 -0.0226 1.0000 1.0000 -1.500 -0.0143 0.02437 0.00926 -0.0216 1.0000 1.0000 -1.250 0.0097 0.02441 0.00897 -0.0207 1.0000 1.0000 -1.000 0.0337 0.02447 0.00878 -0.0198 1.0000 1.0000 -0.750 0.0577 0.02455 0.00870 -0.0189 1.0000 1.0000 -0.500 0.0819 0.02465 0.00869 -0.0179 1.0000 1.0000 -0.250 0.1064 0.02475 0.00887 -0.0170 1.0000 1.0000 0.000 0.1315 0.02487 0.00922 -0.0160 1.0000 1.0000 0.250 0.1586 0.02495 0.00979 -0.0149 1.0000 1.0000 0.500 0.1913 0.02492 0.01064 -0.0143 1.0000 1.0000 0.750 0.2922 0.02828 0.01199 -0.0230 0.4496 1.0000 1.000 0.3196 0.03012 0.01325 -0.0222 0.4169 1.0000 1.250 0.3500 0.03182 0.01467 -0.0219 0.3970 1.0000 1.500 0.3831 0.03346 0.01628 -0.0223 0.3841 1.0000 1.750 0.4163 0.03525 0.01809 -0.0228 0.3735 1.0000 2.000 0.4483 0.03713 0.02005 -0.0232 0.3612 1.0000 2.250 0.4802 0.03896 0.02217 -0.0239 0.3501 1.0000 2.500 0.5111 0.04116 0.02452 -0.0244 0.3427 1.0000 2.750 0.5437 0.04338 0.02718 -0.0256 0.3431 1.0000 3.000 0.5765 0.04565 0.03011 -0.0276 0.3475 1.0000 3.250 0.6076 0.04833 0.03332 -0.0297 0.3537 1.0000 3.500 0.6371 0.05135 0.03677 -0.0315 0.3602 1.0000 3.750 0.6653 0.05457 0.04059 -0.0352 0.3720 1.0000 4.000 0.6922 0.05826 0.04477 -0.0395 0.3856 1.0000 4.250 0.7155 0.06236 0.04928 -0.0444 0.4023 1.0000 4.500 0.7346 0.06686 0.05409 -0.0499 0.4225 1.0000 4.750 0.7498 0.07163 0.05910 -0.0560 0.4465 1.0000 5.000 0.7532 0.07672 0.06433 -0.0632 0.4759 1.0000