XFOIL Version 6.94 Calculated polar for: NACA 67,1-215 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 -0.0347 0.02429 0.02030 0.0019 0.8140 0.8598 0.500 0.0347 0.02396 0.01994 -0.0013 0.8112 0.8619 1.000 0.1090 0.02384 0.01983 -0.0054 0.8090 0.8643 1.500 0.1852 0.02329 0.01927 -0.0096 0.8063 0.8661 2.000 0.2262 0.02348 0.01947 -0.0074 0.7989 0.8689 2.500 0.2077 0.02275 0.01876 0.0065 0.7799 0.8753 3.000 0.2542 0.02157 0.01759 0.0085 0.7708 0.8791 4.500 0.4641 0.01363 0.00902 0.0057 0.6056 0.8869 5.000 0.4261 0.01453 0.00953 0.0234 0.5162 0.8950 5.500 0.4050 0.01554 0.01013 0.0369 0.4186 0.8981 6.000 0.3838 0.01611 0.01037 0.0498 0.3350 0.9011 6.500 0.3828 0.01632 0.01018 0.0585 0.2439 0.9057 7.000 0.3972 0.01697 0.01045 0.0642 0.1575 0.9091 7.500 0.4218 0.01764 0.01089 0.0682 0.0954 0.9121 8.000 0.4452 0.01857 0.01170 0.0722 0.0570 0.9155 8.500 0.4743 0.01935 0.01251 0.0754 0.0380 0.9190 9.000 0.4900 0.02096 0.01429 0.0807 0.0228 0.9237 9.500 0.5186 0.02212 0.01559 0.0837 0.0217 0.9284 10.000 0.5473 0.02352 0.01717 0.0866 0.0208 0.9335 10.500 0.5814 0.02519 0.01905 0.0888 0.0203 0.9390 11.000 0.6259 0.02737 0.02155 0.0897 0.0202 0.9428 11.500 0.6724 0.03062 0.02517 0.0900 0.0204 0.9458 12.000 0.6903 0.03406 0.02896 0.0936 0.0203 0.9490 12.500 0.7035 0.03699 0.03218 0.0962 0.0198 0.9587 13.000 0.7176 0.04119 0.03671 0.0970 0.0193 0.9640 13.500 0.7224 0.04581 0.04164 0.0976 0.0187 0.9684 14.000 0.7114 0.05195 0.04814 0.0981 0.0186 0.9733 14.500 0.6854 0.05982 0.05636 0.0971 0.0185 0.9787 15.000 0.6443 0.07088 0.06777 0.0931 0.0187 0.9840 15.500 0.5735 0.08997 0.08724 0.0818 0.0191 0.9882