XFOIL Version 6.94 Calculated polar for: NASA/LANGLEY NLF(1)-0215F AIRFO 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.5546 0.01567 0.00982 -0.1202 0.6798 0.7890 0.500 0.6048 0.01543 0.00955 -0.1193 0.6735 0.7929 1.000 0.6597 0.01542 0.00944 -0.1194 0.6674 0.7967 1.500 0.7148 0.01552 0.00953 -0.1200 0.6615 0.8005 2.000 0.7719 0.01552 0.00955 -0.1212 0.6546 0.8032 2.500 0.8338 0.01559 0.00956 -0.1233 0.6480 0.8062 3.000 0.8940 0.01575 0.00973 -0.1251 0.6407 0.8076 3.500 0.9523 0.01580 0.00982 -0.1264 0.6313 0.8099 4.000 1.0137 0.01582 0.00978 -0.1282 0.6216 0.8108 4.500 1.0642 0.01564 0.00973 -0.1276 0.6106 0.8119 5.000 1.1192 0.01556 0.00966 -0.1278 0.5994 0.8137 5.500 1.1698 0.01543 0.00964 -0.1272 0.5865 0.8149 6.000 1.2206 0.01538 0.00970 -0.1266 0.5729 0.8158 6.500 1.2710 0.01530 0.00966 -0.1259 0.5574 0.8168 7.000 1.3156 0.01522 0.00969 -0.1241 0.5384 0.8179 7.500 1.3523 0.01522 0.00986 -0.1207 0.5138 0.8202 8.000 1.3818 0.01538 0.01005 -0.1159 0.4802 0.8215 8.500 1.3980 0.01600 0.01058 -0.1090 0.4306 0.8229 9.000 1.4000 0.01756 0.01184 -0.1003 0.3646 0.8244 9.500 1.3945 0.01986 0.01384 -0.0915 0.3035 0.8260 10.000 1.3897 0.02264 0.01643 -0.0842 0.2528 0.8275 10.500 1.3858 0.02594 0.01958 -0.0782 0.2110 0.8290 11.000 1.3847 0.02962 0.02316 -0.0736 0.1768 0.8310 11.500 1.3859 0.03350 0.02699 -0.0700 0.1492 0.8329 12.000 1.3876 0.03775 0.03120 -0.0672 0.1275 0.8346 12.500 1.3948 0.04185 0.03534 -0.0652 0.1096 0.8365 13.000 1.4006 0.04637 0.03990 -0.0635 0.0960 0.8384 13.500 1.4062 0.05119 0.04476 -0.0624 0.0850 0.8405 14.000 1.4144 0.05603 0.04969 -0.0617 0.0755 0.8427 14.500 1.4226 0.06112 0.05489 -0.0614 0.0674 0.8450 15.000 1.4306 0.06651 0.06040 -0.0615 0.0602 0.8473 15.500 1.4379 0.07224 0.06627 -0.0621 0.0536 0.8496 16.000 1.4439 0.07838 0.07254 -0.0631 0.0480 0.8520 16.500 1.4483 0.08488 0.07917 -0.0643 0.0434 0.8549 17.000 1.4532 0.09153 0.08603 -0.0660 0.0393 0.8582 17.500 1.4568 0.09857 0.09330 -0.0681 0.0358 0.8618 18.000 1.4586 0.10595 0.10086 -0.0706 0.0330 0.8657 18.500 1.4602 0.11349 0.10857 -0.0735 0.0307 0.8700 19.000 1.4601 0.12122 0.11658 -0.0767 0.0286 0.8759 19.500 1.4591 0.12905 0.12457 -0.0803 0.0267 0.8842 20.000 1.4550 0.13742 0.13326 -0.0844 0.0249 0.8966 21.000 1.4459 0.15385 0.15017 -0.0938 0.0221 1.0000 21.500 1.4397 0.16337 0.15993 -0.1002 0.0207 1.0000 22.000 1.4362 0.17208 0.16874 -0.1064 0.0194 1.0000 22.500 1.4217 0.18342 0.18042 -0.1146 0.0182 1.0000 23.000 1.4215 0.19119 0.18818 -0.1205 0.0167 1.0000 23.500 1.3961 0.20534 0.20274 -0.1311 0.0158 1.0000 24.000 1.3835 0.21646 0.21402 -0.1397 0.0146 1.0000