XFOIL Version 6.94 Calculated polar for: NACA 63-412 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.3387 0.00762 0.00239 -0.0782 0.7184 0.7115 1.000 0.4509 0.00770 0.00252 -0.0781 0.6876 0.7456 1.500 0.5069 0.00779 0.00263 -0.0780 0.6720 0.7626 2.000 0.5626 0.00791 0.00276 -0.0778 0.6559 0.7805 2.500 0.6179 0.00803 0.00290 -0.0776 0.6381 0.7982 3.000 0.6720 0.00814 0.00307 -0.0770 0.6167 0.8163 3.500 0.7249 0.00826 0.00324 -0.0763 0.5910 0.8338 4.000 0.7758 0.00842 0.00341 -0.0751 0.5521 0.8525 4.500 0.8265 0.00864 0.00367 -0.0739 0.5180 0.8719 5.000 0.8733 0.00901 0.00398 -0.0721 0.4646 0.8945 5.500 0.9068 0.01001 0.00449 -0.0680 0.3353 0.9248 6.000 0.9383 0.01206 0.00567 -0.0648 0.1570 1.0000 6.500 0.9753 0.01412 0.00704 -0.0629 0.0546 1.0000 7.000 1.0183 0.01538 0.00815 -0.0614 0.0335 1.0000 7.500 1.0599 0.01655 0.00931 -0.0596 0.0256 1.0000 8.500 1.1331 0.01899 0.01187 -0.0544 0.0188 1.0000 9.000 1.1577 0.02042 0.01338 -0.0498 0.0173 1.0000 10.000 1.2181 0.02278 0.01590 -0.0436 0.0132 1.0000 10.500 1.2402 0.02460 0.01780 -0.0401 0.0122 1.0000 11.000 1.2622 0.02657 0.01994 -0.0369 0.0114 1.0000 11.500 1.2827 0.02882 0.02234 -0.0341 0.0107 1.0000 12.000 1.3016 0.03136 0.02503 -0.0316 0.0101 1.0000 12.500 1.3169 0.03438 0.02820 -0.0295 0.0096 1.0000 13.000 1.3282 0.03801 0.03202 -0.0276 0.0093 1.0000 13.500 1.3408 0.04174 0.03599 -0.0264 0.0089 1.0000 14.000 1.3548 0.04555 0.04001 -0.0257 0.0082 1.0000 14.500 1.3678 0.04965 0.04429 -0.0255 0.0076 1.0000 15.000 1.3729 0.05498 0.04982 -0.0259 0.0072 1.0000 15.500 1.3733 0.06134 0.05642 -0.0270 0.0069 1.0000 16.000 1.3731 0.06824 0.06361 -0.0288 0.0067 1.0000 16.500 1.3684 0.07633 0.07199 -0.0317 0.0064 1.0000 17.000 1.3597 0.08565 0.08160 -0.0356 0.0061 1.0000 17.500 1.3440 0.09688 0.09314 -0.0411 0.0059 1.0000 18.000 1.3236 0.10980 0.10638 -0.0482 0.0057 1.0000 18.500 1.2921 0.12594 0.12286 -0.0577 0.0058 1.0000 19.000 1.2507 0.14541 0.14267 -0.0700 0.0059 1.0000 19.500 1.1917 0.17059 0.16819 -0.0862 0.0065 1.0000