XFOIL Version 6.94 Calculated polar for: ARA-D BL 13% 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2077 0.04701 0.03948 -0.0137 1.0000 0.3029 -2.750 -0.1811 0.04501 0.03713 -0.0155 1.0000 0.3081 -2.500 -0.1585 0.04320 0.03532 -0.0158 1.0000 0.3158 -2.250 -0.1348 0.04190 0.03397 -0.0162 1.0000 0.3262 -2.000 -0.1078 0.04075 0.03263 -0.0173 1.0000 0.3383 -1.750 -0.0843 0.03995 0.03195 -0.0174 1.0000 0.3532 -1.500 0.1672 0.03115 0.02222 -0.0333 0.5226 0.4888 -1.250 0.1895 0.03106 0.02163 -0.0301 0.4593 0.5439 -1.000 0.2091 0.03029 0.02152 -0.0262 0.4285 0.6507 -0.750 0.2785 0.03027 0.02085 -0.0311 0.3965 1.0000 -0.500 0.3089 0.03143 0.02137 -0.0309 0.3796 1.0000 -0.250 0.3380 0.03265 0.02206 -0.0305 0.3661 1.0000 0.000 0.3673 0.03389 0.02305 -0.0304 0.3553 1.0000 0.250 0.3969 0.03536 0.02412 -0.0302 0.3484 1.0000 0.500 0.4263 0.03695 0.02577 -0.0305 0.3435 1.0000 0.750 0.4547 0.03872 0.02756 -0.0308 0.3395 1.0000 1.000 0.4820 0.04063 0.02947 -0.0310 0.3358 1.0000 1.250 0.5083 0.04259 0.03134 -0.0310 0.3317 1.0000 1.500 0.5342 0.04474 0.03324 -0.0307 0.3273 1.0000 1.750 0.5568 0.04733 0.03604 -0.0313 0.3249 1.0000 2.000 0.5778 0.05027 0.03916 -0.0320 0.3243 1.0000 2.250 0.5977 0.05351 0.04251 -0.0327 0.3254 1.0000 2.500 0.6169 0.05693 0.04597 -0.0332 0.3271 1.0000 2.750 0.6039 0.06446 0.05428 -0.0378 0.3391 1.0000 3.000 0.6116 0.06934 0.05918 -0.0394 0.3455 1.0000 3.500 0.4893 0.09311 0.08360 -0.0581 0.4310 1.0000 4.250 0.3496 0.11119 0.10202 -0.0758 0.7186 1.0000 4.750 0.3741 0.11703 0.10758 -0.0758 0.6905 1.0000 5.000 0.4072 0.12254 0.11292 -0.0782 0.6768 1.0000