XFOIL Version 6.94 Calculated polar for: NACA 66-018 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 0.0000 0.01429 0.00945 0.0000 0.7707 0.7707 0.500 0.0501 0.01415 0.00928 0.0001 0.7673 0.7750 1.000 0.1051 0.01395 0.00906 -0.0005 0.7645 0.7773 1.500 0.1605 0.01390 0.00905 -0.0010 0.7606 0.7790 2.000 0.2115 0.01395 0.00917 -0.0009 0.7548 0.7809 2.500 0.2723 0.01357 0.00881 -0.0020 0.7506 0.7833 3.000 0.3272 0.01318 0.00847 -0.0022 0.7417 0.7857 3.500 0.3890 0.01230 0.00756 -0.0031 0.7302 0.7881 4.000 0.4514 0.01106 0.00621 -0.0037 0.7084 0.7903 4.500 0.5013 0.01060 0.00588 -0.0025 0.6857 0.7925 5.000 0.5485 0.01030 0.00567 -0.0009 0.6505 0.7950 5.500 0.5551 0.01089 0.00559 0.0084 0.4953 0.7989 6.000 0.5241 0.01246 0.00653 0.0234 0.3682 0.8043 6.500 0.5070 0.01478 0.00818 0.0341 0.2224 0.8086 7.000 0.5024 0.01724 0.00983 0.0419 0.0762 0.8138 7.500 0.5273 0.01861 0.01102 0.0454 0.0499 0.8185 8.000 0.5556 0.01983 0.01226 0.0485 0.0422 0.8224 8.500 0.5835 0.02118 0.01363 0.0514 0.0377 0.8273 9.000 0.6146 0.02242 0.01492 0.0537 0.0346 0.8325 9.500 0.6415 0.02390 0.01648 0.0565 0.0320 0.8374 10.000 0.6719 0.02534 0.01797 0.0587 0.0303 0.8434 10.500 0.7027 0.02697 0.01961 0.0609 0.0289 0.8494 11.000 0.7387 0.02839 0.02119 0.0625 0.0277 0.8559 11.500 0.7758 0.02984 0.02272 0.0638 0.0264 0.8626 12.000 0.8228 0.03151 0.02443 0.0641 0.0253 0.8691 12.500 0.8620 0.03350 0.02671 0.0652 0.0248 0.8767 13.000 0.8949 0.03585 0.02938 0.0668 0.0240 0.8855 13.500 0.9236 0.03878 0.03264 0.0686 0.0237 0.8946 14.000 0.9493 0.04109 0.03506 0.0703 0.0228 0.9061 14.500 0.9754 0.04468 0.03886 0.0717 0.0223 0.9190 15.000 0.9761 0.04971 0.04434 0.0746 0.0223 0.9369 15.500 0.9570 0.05644 0.05170 0.0748 0.0222 0.9593 16.500 0.9183 0.07352 0.06957 0.0705 0.0223 1.0000