XFOIL Version 6.94 Calculated polar for: nacc632 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 -0.0206 0.03182 0.01537 -0.0236 0.9994 1.0006 -2.750 0.0004 0.03216 0.01544 -0.0232 0.9994 1.0006 -2.500 0.0213 0.03256 0.01559 -0.0228 0.9994 1.0006 -2.250 0.0420 0.03303 0.01585 -0.0224 0.9994 1.0006 -2.000 0.0626 0.03357 0.01620 -0.0219 0.9994 1.0006 -1.750 0.0831 0.03416 0.01663 -0.0215 0.9994 1.0006 -1.500 0.1035 0.03481 0.01712 -0.0212 0.9994 1.0006 -1.250 0.1240 0.03551 0.01771 -0.0209 0.9994 1.0006 -1.000 0.1445 0.03627 0.01837 -0.0207 0.9994 1.0006 -0.750 0.1648 0.03708 0.01912 -0.0205 0.9994 1.0006 -0.500 0.1850 0.03795 0.01993 -0.0204 0.9994 1.0006 -0.250 0.2049 0.03888 0.02084 -0.0203 0.9994 1.0006 0.000 0.2246 0.03987 0.02184 -0.0203 0.9994 1.0006 0.250 0.2439 0.04094 0.02294 -0.0203 0.9994 1.0006 0.500 0.2628 0.04211 0.02416 -0.0205 0.9994 1.0006 0.750 0.2811 0.04339 0.02553 -0.0207 0.9994 1.0006 1.000 0.2986 0.04483 0.02709 -0.0211 0.9994 1.0006 1.250 0.3147 0.04648 0.02889 -0.0216 0.9994 1.0006 1.500 0.3288 0.04847 0.03106 -0.0224 0.9994 1.0006 1.750 0.3390 0.05105 0.03386 -0.0237 0.9994 1.0006 2.000 0.3429 0.05465 0.03768 -0.0259 0.9994 1.0006 2.250 0.3426 0.05906 0.04226 -0.0289 0.9994 1.0006 2.500 0.5434 0.06350 0.04739 -0.0660 0.8452 1.0006 2.750 0.6213 0.06526 0.04948 -0.0742 0.7722 1.0006 3.000 0.6658 0.06733 0.05170 -0.0767 0.7250 1.0006 3.250 0.6925 0.07002 0.05449 -0.0772 0.6922 1.0006 3.500 0.7216 0.07255 0.05716 -0.0776 0.6631 1.0006 3.750 0.7284 0.07620 0.06086 -0.0770 0.6417 1.0006 4.000 0.7373 0.08019 0.06490 -0.0772 0.6265 1.0006 4.250 0.7494 0.08419 0.06895 -0.0776 0.6132 1.0006 4.500 0.7364 0.08914 0.07389 -0.0769 0.6065 1.0006 4.750 0.7427 0.09317 0.07795 -0.0768 0.5941 1.0006 5.000 0.7351 0.09801 0.08278 -0.0766 0.5902 1.0006