XFOIL Version 6.96 Calculated polar for: NACA 66-206 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 ------ -------- --------- --------- -------- -------- -------- -9.250 -0.5252 0.10135 0.09898 -0.0124 1.0000 0.0047 -9.000 -0.5221 0.09756 0.09521 -0.0141 1.0000 0.0047 -8.750 -0.5201 0.09354 0.09121 -0.0159 1.0000 0.0047 -8.500 -0.5187 0.08954 0.08724 -0.0178 1.0000 0.0047 -8.250 -0.5170 0.08590 0.08362 -0.0198 1.0000 0.0047 -8.000 -0.5176 0.08203 0.07978 -0.0220 1.0000 0.0047 -7.750 -0.5214 0.07811 0.07590 -0.0250 1.0000 0.0047 -7.500 -0.5194 0.07378 0.07155 -0.0285 1.0000 0.0047 -7.250 -0.5227 0.06883 0.06658 -0.0307 1.0000 0.0044 -7.000 -0.5230 0.06412 0.06184 -0.0326 1.0000 0.0041 -6.750 -0.5184 0.06007 0.05773 -0.0337 1.0000 0.0039 -6.500 -0.5136 0.05597 0.05355 -0.0341 1.0000 0.0037 -6.250 -0.4995 0.05131 0.04876 -0.0361 0.9978 0.0035 -6.000 -0.4746 0.04572 0.04295 -0.0397 0.9924 0.0033 -5.750 -0.4475 0.04020 0.03716 -0.0426 0.9880 0.0031 -5.500 -0.4201 0.03482 0.03145 -0.0444 0.9833 0.0030 -5.250 -0.3912 0.02967 0.02591 -0.0457 0.9792 0.0029 -5.000 -0.3627 0.02490 0.02069 -0.0462 0.9747 0.0030 -4.750 -0.3337 0.02034 0.01562 -0.0461 0.9702 0.0034 -4.250 -0.2758 0.01456 0.00902 -0.0461 0.9618 0.0053 -4.000 -0.2467 0.01317 0.00741 -0.0464 0.9577 0.0059 -3.750 -0.2179 0.01185 0.00591 -0.0466 0.9539 0.0065 -3.500 -0.1911 0.01094 0.00489 -0.0463 0.9483 0.0084 -3.250 -0.1627 0.01053 0.00435 -0.0465 0.9437 0.0103 -3.000 -0.1369 0.00958 0.00331 -0.0463 0.9383 0.0135 -2.750 -0.1100 0.00917 0.00282 -0.0462 0.9329 0.0157 -2.500 -0.0824 0.00887 0.00246 -0.0462 0.9286 0.0184 -2.250 -0.0558 0.00854 0.00205 -0.0460 0.9228 0.0201 -2.000 -0.0286 0.00834 0.00174 -0.0459 0.9178 0.0221 -1.750 -0.0016 0.00807 0.00144 -0.0458 0.9128 0.0330 -1.500 0.0252 0.00780 0.00128 -0.0458 0.9077 0.0745 -1.250 0.0466 0.00605 0.00105 -0.0458 0.9028 0.5418 -1.000 0.0666 0.00533 0.00118 -0.0441 0.8969 0.7884 -0.750 0.0905 0.00527 0.00123 -0.0429 0.8920 0.8392 -0.500 0.1146 0.00526 0.00128 -0.0419 0.8876 0.8688 -0.250 0.1398 0.00526 0.00130 -0.0411 0.8824 0.8852 0.000 0.1666 0.00527 0.00129 -0.0409 0.8777 0.8921 0.250 0.1934 0.00527 0.00129 -0.0407 0.8716 0.8976 0.500 0.2199 0.00528 0.00127 -0.0403 0.8616 0.9031 0.750 0.2448 0.00527 0.00123 -0.0395 0.8419 0.9085 1.000 0.2705 0.00530 0.00121 -0.0389 0.8235 0.9147 1.250 0.2961 0.00532 0.00122 -0.0383 0.8052 0.9200 1.500 0.3206 0.00540 0.00120 -0.0374 0.7697 0.9265 1.750 0.3346 0.00625 0.00120 -0.0343 0.5491 0.9346 2.000 0.3480 0.00784 0.00158 -0.0323 0.2352 0.9443 2.250 0.3691 0.00873 0.00187 -0.0314 0.0669 0.9532 2.500 0.3958 0.00905 0.00209 -0.0314 0.0353 0.9613 2.750 0.4244 0.00933 0.00239 -0.0317 0.0260 0.9697 3.000 0.4550 0.00957 0.00268 -0.0325 0.0212 0.9783 3.250 0.4862 0.00995 0.00315 -0.0334 0.0177 0.9888 3.500 0.5125 0.01076 0.00407 -0.0334 0.0145 1.0000 3.750 0.5376 0.01124 0.00462 -0.0330 0.0136 1.0000 4.000 0.5635 0.01158 0.00501 -0.0328 0.0116 1.0000 4.250 0.5882 0.01224 0.00576 -0.0323 0.0098 1.0000 4.500 0.6130 0.01285 0.00638 -0.0320 0.0067 1.0000 4.750 0.6378 0.01360 0.00728 -0.0315 0.0057 1.0000 5.000 0.6622 0.01472 0.00856 -0.0308 0.0044 1.0000 5.250 0.6865 0.01623 0.01027 -0.0301 0.0038 1.0000 5.500 0.7107 0.01830 0.01264 -0.0292 0.0035 1.0000 5.750 0.7330 0.02206 0.01689 -0.0276 0.0034 1.0000 6.000 0.7513 0.02779 0.02324 -0.0249 0.0035 1.0000 6.250 0.7667 0.03374 0.02972 -0.0221 0.0037 1.0000 6.500 0.7803 0.03916 0.03556 -0.0199 0.0040 1.0000 6.750 0.7920 0.04432 0.04106 -0.0180 0.0042 1.0000 7.000 0.8016 0.04939 0.04641 -0.0165 0.0045 1.0000 7.250 0.8099 0.05417 0.05142 -0.0155 0.0045 1.0000 7.500 0.8159 0.05897 0.05642 -0.0147 0.0044 1.0000 7.750 0.8192 0.06376 0.06137 -0.0144 0.0042 1.0000 8.000 0.8256 0.06677 0.06449 -0.0145 0.0035 1.0000 8.250 0.8297 0.06930 0.06709 -0.0147 0.0031 1.0000 8.500 0.8226 0.07344 0.07133 -0.0154 0.0028 1.0000 8.750 0.8134 0.07785 0.07580 -0.0155 0.0030 1.0000 9.000 0.7959 0.08290 0.08089 -0.0182 0.0028 1.0000