XFOIL Version 6.96 Calculated polar for: Coanda 1 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.000 -0.3487 0.07964 0.07647 -0.0198 1.0000 0.0465 -7.750 -0.3491 0.07648 0.07335 -0.0195 1.0000 0.0475 -7.500 -0.4444 0.08629 0.08304 -0.0186 1.0000 0.0451 -7.250 -0.4393 0.08311 0.07988 -0.0169 1.0000 0.0460 -7.000 -0.4358 0.08029 0.07708 -0.0165 1.0000 0.0470 -6.750 -0.4325 0.07737 0.07418 -0.0166 1.0000 0.0482 -6.500 -0.4283 0.07442 0.07124 -0.0171 1.0000 0.0495 -6.250 -0.4229 0.07131 0.06812 -0.0180 1.0000 0.0515 -6.000 -0.4071 0.06890 0.06552 -0.0243 1.0000 0.0544 -5.750 -0.3970 0.06586 0.06233 -0.0248 1.0000 0.0546 -5.500 -0.3879 0.06238 0.05874 -0.0240 1.0000 0.0547 -5.250 -0.3786 0.05888 0.05514 -0.0228 1.0000 0.0547 -5.000 -0.3694 0.05513 0.05129 -0.0214 1.0000 0.0547 -4.750 -0.3768 0.04349 0.03958 -0.0190 1.0000 0.0346 -4.500 -0.3664 0.04052 0.03649 -0.0168 1.0000 0.0314 -4.250 -0.3566 0.03602 0.03173 -0.0145 1.0000 0.0292 -4.000 -0.3567 0.02275 0.01711 -0.0089 1.0000 0.0243 -3.750 -0.3397 0.01942 0.01310 -0.0063 1.0000 0.0242 -3.500 -0.3197 0.01729 0.01048 -0.0044 1.0000 0.0257 -3.250 -0.2988 0.01641 0.00941 -0.0029 1.0000 0.0298 -3.000 -0.2765 0.01538 0.00809 -0.0013 1.0000 0.0317 -2.750 -0.2547 0.01482 0.00739 0.0003 1.0000 0.0362 -2.500 -0.2343 0.01587 0.00873 0.0018 0.9997 0.0491 -2.250 -0.1878 0.01867 0.01149 -0.0023 0.9915 0.0824 -2.000 -0.1478 0.01923 0.01205 -0.0056 0.9848 0.0972 -1.750 -0.1117 0.01954 0.01243 -0.0085 0.9778 0.1252 -1.250 -0.0351 0.01884 0.01136 -0.0131 0.9647 0.1433 -1.000 0.0056 0.01753 0.00999 -0.0157 0.9590 0.1383 -0.750 0.0482 0.01693 0.00910 -0.0182 0.9513 0.1322 -0.500 0.0948 0.01650 0.00856 -0.0217 0.9440 0.1308 -0.250 0.1369 0.01552 0.00757 -0.0243 0.9342 0.1318 0.000 0.1848 0.01492 0.00694 -0.0279 0.9260 0.1306 0.250 0.2313 0.01407 0.00610 -0.0313 0.9172 0.1306 0.500 0.2746 0.01330 0.00538 -0.0340 0.9048 0.1308 1.000 0.3533 0.01201 0.00430 -0.0382 0.8757 0.1339 1.250 0.3879 0.01161 0.00397 -0.0393 0.8541 0.1356 1.500 0.4218 0.01129 0.00368 -0.0402 0.8282 0.1375 1.750 0.4533 0.01108 0.00347 -0.0406 0.7975 0.1406 2.000 0.4830 0.01099 0.00331 -0.0406 0.7654 0.1440 2.250 0.5092 0.01094 0.00323 -0.0399 0.7319 0.1473 2.500 0.5341 0.01102 0.00321 -0.0390 0.6980 0.1507 2.750 0.5559 0.01119 0.00321 -0.0374 0.6505 0.1558 3.000 0.5736 0.01147 0.00320 -0.0349 0.5789 0.1638 3.250 0.7215 0.01184 0.00407 -0.0630 0.3156 1.0000 3.500 0.7410 0.01246 0.00431 -0.0616 0.2530 1.0000 3.750 0.7622 0.01296 0.00459 -0.0603 0.2086 1.0000 4.250 0.8050 0.01396 0.00533 -0.0578 0.1394 1.0000 4.500 0.8199 0.01531 0.00601 -0.0557 0.0318 1.0000 4.750 0.8415 0.01588 0.00658 -0.0543 0.0265 1.0000 5.000 0.8633 0.01645 0.00729 -0.0528 0.0246 1.0000 5.250 0.8845 0.01710 0.00814 -0.0512 0.0242 1.0000 5.500 0.9048 0.01784 0.00908 -0.0495 0.0242 1.0000 5.750 0.9232 0.01877 0.01020 -0.0475 0.0243 1.0000 6.000 0.9393 0.01992 0.01153 -0.0451 0.0241 1.0000 6.250 0.9555 0.02111 0.01281 -0.0428 0.0235 1.0000 6.500 0.9718 0.02253 0.01431 -0.0405 0.0230 1.0000 6.750 0.9894 0.02443 0.01626 -0.0383 0.0239 1.0000 7.000 1.0103 0.02658 0.01847 -0.0367 0.0250 1.0000 7.250 1.0346 0.02752 0.01969 -0.0349 0.0280 1.0000 7.500 1.0562 0.03131 0.02385 -0.0330 0.0326 1.0000 7.750 1.0777 0.03418 0.02736 -0.0299 0.0409 1.0000 10.500 0.7594 0.09388 0.09083 -0.0079 0.0869 1.0000 10.750 0.7392 0.10160 0.09850 -0.0109 0.0870 1.0000