XFOIL Version 6.96 Calculated polar for: NACA 63-206 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -7.750 -0.5595 0.08681 0.08215 -0.0092 1.0000 0.0770 -7.500 -0.5617 0.08245 0.07784 -0.0167 1.0000 0.0795 -7.250 -0.5652 0.07844 0.07361 -0.0278 1.0000 0.0809 -7.000 -0.5511 0.07336 0.06869 -0.0216 1.0000 0.0851 -6.750 -0.5426 0.06941 0.06470 -0.0244 1.0000 0.0908 -6.500 -0.5361 0.06428 0.05939 -0.0302 1.0000 0.0962 -6.250 -0.5229 0.06089 0.05603 -0.0295 1.0000 0.1034 -6.000 -0.5111 0.05659 0.05161 -0.0317 1.0000 0.1121 -5.750 -0.4971 0.05280 0.04768 -0.0334 1.0000 0.1248 -5.500 -0.4819 0.04941 0.04422 -0.0339 1.0000 0.1401 -5.250 -0.4665 0.04613 0.04078 -0.0349 1.0000 0.1641 -5.000 -0.4513 0.04327 0.03793 -0.0343 1.0000 0.1919 -4.250 -0.3561 0.03103 0.02355 -0.0384 1.0000 0.1206 -4.000 -0.3224 0.02582 0.01764 -0.0371 1.0000 0.0733 -3.750 -0.2939 0.02329 0.01455 -0.0359 1.0000 0.0640 -3.500 -0.2673 0.02097 0.01181 -0.0349 1.0000 0.0608 -3.250 -0.2415 0.01919 0.00981 -0.0338 1.0000 0.0616 -3.000 -0.2164 0.01798 0.00839 -0.0328 1.0000 0.0688 -2.750 -0.1928 0.01631 0.00684 -0.0317 1.0000 0.0727 -2.500 -0.1691 0.01520 0.00572 -0.0306 1.0000 0.0798 -2.250 -0.1451 0.01403 0.00462 -0.0299 1.0000 0.1000 -2.000 -0.1389 0.01063 0.00424 -0.0250 1.0000 0.7596 -1.750 -0.1021 0.01022 0.00390 -0.0226 1.0000 1.0000 -1.500 -0.0853 0.01013 0.00356 -0.0215 1.0000 1.0000 -1.250 -0.0639 0.01011 0.00331 -0.0210 1.0000 1.0000 -1.000 -0.0413 0.01013 0.00315 -0.0207 1.0000 1.0000 -0.750 -0.0182 0.01019 0.00304 -0.0205 1.0000 1.0000 -0.500 0.0048 0.01028 0.00300 -0.0202 1.0000 1.0000 -0.250 0.0276 0.01041 0.00299 -0.0199 1.0000 1.0000 0.000 0.0502 0.01056 0.00306 -0.0197 1.0000 1.0000 0.250 0.0725 0.01075 0.00318 -0.0194 1.0000 1.0000 0.500 0.0944 0.01098 0.00336 -0.0191 1.0000 1.0000 0.750 0.1160 0.01124 0.00360 -0.0188 1.0000 1.0000 1.000 0.1373 0.01154 0.00389 -0.0186 1.0000 1.0000 1.250 0.1581 0.01188 0.00426 -0.0184 1.0000 1.0000 1.500 0.1829 0.01228 0.00469 -0.0191 0.9981 1.0000 1.750 0.2319 0.01278 0.00530 -0.0242 0.9858 1.0000 2.000 0.2807 0.01322 0.00589 -0.0292 0.9731 1.0000 2.250 0.3307 0.01357 0.00651 -0.0342 0.9592 1.0000 2.500 0.3837 0.01378 0.00699 -0.0393 0.9440 1.0000 2.750 0.4717 0.01240 0.00613 -0.0464 0.8905 1.0000 3.000 0.4957 0.01151 0.00525 -0.0406 0.7976 1.0000 3.250 0.4980 0.01263 0.00462 -0.0316 0.3809 1.0000 3.500 0.5105 0.01607 0.00628 -0.0295 0.0905 1.0000 3.750 0.5331 0.01750 0.00773 -0.0283 0.0775 1.0000 4.000 0.5564 0.01901 0.00916 -0.0273 0.0656 1.0000 4.250 0.5821 0.02071 0.01084 -0.0265 0.0601 1.0000 4.500 0.6093 0.02319 0.01331 -0.0259 0.0580 1.0000 4.750 0.6379 0.02555 0.01593 -0.0251 0.0583 1.0000 5.000 0.6659 0.02807 0.01884 -0.0241 0.0589 1.0000 5.250 0.6900 0.03159 0.02286 -0.0233 0.0564 1.0000 5.500 0.7181 0.03488 0.02699 -0.0214 0.0671 1.0000 8.250 0.7185 0.08362 0.07965 -0.0285 0.1186 1.0000 8.500 0.7072 0.08903 0.08502 -0.0316 0.1147 1.0000 8.750 0.7212 0.09375 0.08975 -0.0290 0.1084 1.0000 9.000 0.7001 0.09958 0.09551 -0.0346 0.1071 1.0000 9.250 0.6885 0.10486 0.10076 -0.0383 0.1034 1.0000 9.500 0.7148 0.10941 0.10534 -0.0334 0.0958 1.0000 9.750 0.6969 0.11464 0.11052 -0.0386 0.0952 1.0000 10.000 0.6823 0.11943 0.11526 -0.0432 0.0931 1.0000