XFOIL Version 6.94 Calculated polar for: FX 38-153 AIRFOIL 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 ------- -------- --------- --------- -------- ------- ------- 1.000 0.1329 0.02091 0.01603 -0.0138 0.8123 0.8506 1.500 0.2018 0.01979 0.01488 -0.0157 0.8067 0.8531 2.000 0.2773 0.01862 0.01370 -0.0190 0.8039 0.8555 4.000 0.5418 0.01473 0.00983 -0.0272 0.7386 0.8711 4.500 0.5954 0.01419 0.00923 -0.0267 0.6971 0.8735 5.000 0.6446 0.01407 0.00888 -0.0255 0.6362 0.8766 5.500 0.6749 0.01450 0.00895 -0.0212 0.5604 0.8807 6.000 0.7014 0.01520 0.00933 -0.0170 0.4887 0.8844 6.500 0.7323 0.01612 0.00996 -0.0141 0.4197 0.8884 7.000 0.7649 0.01689 0.01055 -0.0115 0.3645 0.8910 7.500 0.7912 0.01769 0.01116 -0.0075 0.3076 0.8938 8.000 0.8181 0.01865 0.01195 -0.0038 0.2581 0.8975 8.500 0.8516 0.01969 0.01286 -0.0017 0.2157 0.9004 9.000 0.8856 0.02092 0.01390 0.0000 0.1706 0.9030 9.500 0.9225 0.02213 0.01503 0.0010 0.1385 0.9055 10.000 0.9574 0.02361 0.01635 0.0021 0.0998 0.9078 10.500 0.9917 0.02520 0.01787 0.0031 0.0710 0.9098 11.000 1.0148 0.02686 0.01946 0.0061 0.0494 0.9128 11.500 1.0343 0.02881 0.02138 0.0093 0.0319 0.9165 12.000 1.0561 0.03103 0.02371 0.0117 0.0248 0.9195 12.500 1.0832 0.03304 0.02592 0.0130 0.0209 0.9224 13.000 1.1091 0.03533 0.02838 0.0141 0.0183 0.9254 13.500 1.1321 0.03802 0.03122 0.0150 0.0163 0.9280 14.000 1.1472 0.04123 0.03459 0.0163 0.0145 0.9312 14.500 1.1546 0.04469 0.03830 0.0185 0.0134 0.9361 15.000 1.1583 0.04924 0.04309 0.0195 0.0126 0.9403 15.500 1.1625 0.05434 0.04847 0.0194 0.0120 0.9443 16.000 1.1643 0.06031 0.05474 0.0183 0.0110 0.9479 16.500 1.1534 0.06789 0.06261 0.0166 0.0102 0.9525 17.000 1.1428 0.07574 0.07078 0.0143 0.0096 0.9581 17.500 1.1332 0.08430 0.07968 0.0109 0.0088 0.9639