XFOIL Version 6.96 Calculated polar for: NASA LRN 1015 (NASA TM 102840) 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -10.750 -0.0847 0.08550 0.08170 -0.1066 0.7381 0.0104 -10.500 -0.0841 0.08170 0.07793 -0.1083 0.7368 0.0101 -10.250 -0.0880 0.07704 0.07329 -0.1104 0.7354 0.0103 -10.000 -0.0923 0.07247 0.06875 -0.1126 0.7340 0.0102 -9.750 -0.1010 0.06714 0.06345 -0.1155 0.7325 0.0098 -9.500 -0.1165 0.06021 0.05655 -0.1205 0.7309 0.0099 -9.250 -0.1519 0.05202 0.04828 -0.1258 0.7289 0.0098 -9.000 -0.1847 0.04742 0.04358 -0.1239 0.7269 0.0096 -8.750 -0.2008 0.04317 0.03917 -0.1222 0.7250 0.0096 -8.500 -0.2117 0.03873 0.03450 -0.1200 0.7232 0.0097 -8.250 -0.2165 0.03443 0.02990 -0.1177 0.7215 0.0098 -8.000 -0.2168 0.03017 0.02527 -0.1153 0.7199 0.0099 -7.750 -0.2111 0.02645 0.02113 -0.1130 0.7184 0.0102 -7.500 -0.2008 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1.1205 0.01646 0.01023 -0.0641 0.2833 1.0000 7.500 1.1300 0.01742 0.01112 -0.0623 0.2689 1.0000 7.750 1.1385 0.01844 0.01207 -0.0604 0.2530 1.0000 8.000 1.1477 0.01945 0.01301 -0.0586 0.2380 1.0000 8.250 1.1580 0.02041 0.01392 -0.0570 0.2253 1.0000 8.500 1.1686 0.02137 0.01483 -0.0555 0.2142 1.0000 8.750 1.1802 0.02229 0.01572 -0.0541 0.2032 1.0000 9.000 1.1921 0.02321 0.01661 -0.0528 0.1934 1.0000 9.250 1.2026 0.02422 0.01758 -0.0514 0.1837 1.0000 9.500 1.2148 0.02515 0.01849 -0.0502 0.1744 1.0000 9.750 1.2260 0.02615 0.01946 -0.0489 0.1654 1.0000 10.000 1.2373 0.02716 0.02046 -0.0477 0.1575 1.0000 10.250 1.2499 0.02810 0.02140 -0.0466 0.1510 1.0000 10.500 1.2615 0.02912 0.02242 -0.0455 0.1441 1.0000 10.750 1.2734 0.03016 0.02345 -0.0445 0.1382 1.0000 11.000 1.2858 0.03119 0.02449 -0.0436 0.1321 1.0000 11.250 1.2970 0.03233 0.02563 -0.0426 0.1265 1.0000 11.500 1.3096 0.03339 0.02672 -0.0418 0.1215 1.0000 11.750 1.3206 0.03459 0.02791 -0.0409 0.1158 1.0000 12.000 1.3323 0.03575 0.02910 -0.0401 0.1111 1.0000 12.250 1.3436 0.03698 0.03034 -0.0393 0.1063 1.0000 12.500 1.3540 0.03829 0.03166 -0.0384 0.1020 1.0000 12.750 1.3658 0.03951 0.03292 -0.0378 0.0987 1.0000 13.000 1.3771 0.04079 0.03425 -0.0371 0.0956 1.0000 13.250 1.3873 0.04218 0.03566 -0.0364 0.0921 1.0000 13.500 1.3962 0.04372 0.03722 -0.0357 0.0886 1.0000 13.750 1.4073 0.04507 0.03863 -0.0352 0.0860 1.0000 14.000 1.4176 0.04652 0.04013 -0.0346 0.0834 1.0000 14.250 1.4250 0.04824 0.04186 -0.0340 0.0796 1.0000 14.500 1.4337 0.04988 0.04354 -0.0335 0.0764 1.0000 14.750 1.4429 0.05150 0.04521 -0.0331 0.0733 1.0000 15.000 1.4501 0.05333 0.04707 -0.0326 0.0703 1.0000 15.250 1.4564 0.05529 0.04905 -0.0322 0.0675 1.0000 15.500 1.4647 0.05707 0.05092 -0.0319 0.0662 1.0000 15.750 1.4733 0.05884 0.05277 -0.0317 0.0643 1.0000 16.000 1.4800 0.06084 0.05483 -0.0314 0.0622 1.0000 16.250 1.4854 0.06302 0.05704 -0.0313 0.0599 1.0000 16.500 1.4902 0.06530 0.05939 -0.0311 0.0585 1.0000 16.750 1.4962 0.06746 0.06162 -0.0311 0.0563 1.0000 17.000 1.5020 0.06968 0.06392 -0.0311 0.0541 1.0000 17.250 1.5065 0.07208 0.06639 -0.0312 0.0523 1.0000 17.500 1.5090 0.07478 0.06915 -0.0314 0.0505 1.0000 17.750 1.5113 0.07753 0.07197 -0.0317 0.0487 1.0000 18.000 1.5156 0.08008 0.07462 -0.0320 0.0470 1.0000 18.250 1.5178 0.08293 0.07753 -0.0324 0.0444 1.0000 18.500 1.5165 0.08627 0.08092 -0.0330 0.0418 1.0000 18.750 1.5173 0.08940 0.08414 -0.0337 0.0395 1.0000 19.000 1.5159 0.09287 0.08767 -0.0345 0.0363 1.0000 19.250 1.5129 0.09661 0.09147 -0.0354 0.0331 1.0000