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 = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.5873 0.01259 0.00723 -0.1295 0.6518 0.7555 0.500 0.6455 0.01268 0.00721 -0.1304 0.6460 0.7594 1.000 0.7048 0.01272 0.00721 -0.1317 0.6405 0.7632 1.500 0.7649 0.01271 0.00718 -0.1333 0.6346 0.7662 2.000 0.8264 0.01281 0.00721 -0.1351 0.6285 0.7686 2.500 0.8851 0.01281 0.00717 -0.1363 0.6217 0.7702 3.000 0.9373 0.01267 0.00712 -0.1359 0.6137 0.7722 3.500 0.9935 0.01264 0.00704 -0.1363 0.6042 0.7735 4.000 1.0472 0.01264 0.00712 -0.1364 0.5949 0.7747 4.500 1.1001 0.01262 0.00710 -0.1360 0.5840 0.7766 5.000 1.1524 0.01263 0.00721 -0.1358 0.5730 0.7778 5.500 1.2054 0.01268 0.00724 -0.1356 0.5603 0.7789 6.000 1.2531 0.01262 0.00734 -0.1343 0.5456 0.7801 6.500 1.3000 0.01267 0.00746 -0.1329 0.5275 0.7813 7.000 1.3420 0.01281 0.00764 -0.1305 0.5028 0.7833 7.500 1.3767 0.01311 0.00792 -0.1267 0.4663 0.7846 8.000 1.3938 0.01389 0.00850 -0.1198 0.4112 0.7859 8.500 1.4000 0.01545 0.00972 -0.1114 0.3422 0.7872 9.000 1.4035 0.01736 0.01134 -0.1034 0.2824 0.7884 9.500 1.4078 0.01947 0.01325 -0.0963 0.2332 0.7894 10.000 1.4116 0.02187 0.01551 -0.0899 0.1939 0.7908 10.500 1.4150 0.02465 0.01819 -0.0844 0.1612 0.7925 11.000 1.4195 0.02777 0.02126 -0.0799 0.1348 0.7946 11.500 1.4255 0.03119 0.02464 -0.0762 0.1142 0.7962 12.000 1.4339 0.03473 0.02821 -0.0734 0.0975 0.7977 12.500 1.4413 0.03865 0.03214 -0.0710 0.0845 0.7994 13.000 1.4522 0.04250 0.03607 -0.0692 0.0734 0.8011 13.500 1.4622 0.04666 0.04029 -0.0678 0.0639 0.8029 14.000 1.4715 0.05111 0.04482 -0.0668 0.0564 0.8047 14.500 1.4803 0.05587 0.04967 -0.0661 0.0498 0.8064 15.000 1.4884 0.06093 0.05486 -0.0658 0.0445 0.8081 15.500 1.4953 0.06639 0.06045 -0.0658 0.0397 0.8097 16.000 1.5013 0.07218 0.06640 -0.0663 0.0356 0.8116 16.500 1.5060 0.07838 0.07277 -0.0671 0.0319 0.8139 17.000 1.5074 0.08529 0.07985 -0.0685 0.0291 0.8162 17.500 1.5077 0.09259 0.08733 -0.0703 0.0268 0.8185 18.000 1.5070 0.10030 0.09527 -0.0726 0.0249 0.8210 18.500 1.5039 0.10859 0.10373 -0.0755 0.0231 0.8235 19.000 1.5026 0.11675 0.11213 -0.0788 0.0214 0.8263 19.500 1.4973 0.12570 0.12127 -0.0829 0.0201 0.8291 20.000 1.4907 0.13493 0.13074 -0.0875 0.0189 0.8324 20.500 1.4853 0.14397 0.14003 -0.0924 0.0176 0.8370 21.000 1.4746 0.15399 0.15022 -0.0983 0.0166 0.8419 21.500 1.4677 0.16341 0.15993 -0.1043 0.0154 0.8477 22.000 1.4572 0.17349 0.17020 -0.1110 0.0144 0.8546 22.500 1.4449 0.18383 0.18079 -0.1180 0.0133 0.8656 23.000 1.4312 0.19401 0.19124 -0.1247 0.0122 0.8956 23.500 1.4150 0.20443 0.20186 -0.1322 0.0112 1.0000 24.000 1.4016 0.21559 0.21319 -0.1406 0.0100 1.0000 24.500 1.3883 0.22661 0.22431 -0.1491 0.0091 1.0000 25.000 1.3764 0.23749 0.23535 -0.1574 0.0082 1.0000 25.500 1.3697 0.24699 0.24489 -0.1649 0.0075 1.0000 26.000 1.3599 0.25746 0.25551 -0.1729 0.0071 1.0000 26.500 1.3413 0.27108 0.26934 -0.1827 0.0068 1.0000