XFOIL Version 6.96 Calculated polar for: Coanda 2 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 ------ -------- --------- --------- -------- -------- -------- -8.750 -0.4287 0.09760 0.09421 -0.0187 1.0000 0.0497 -8.500 -0.4410 0.09602 0.09271 -0.0203 1.0000 0.0500 -8.250 -0.4455 0.09357 0.09031 -0.0230 1.0000 0.0502 -8.000 -0.4444 0.09034 0.08710 -0.0259 1.0000 0.0503 -7.750 -0.4436 0.08545 0.08226 -0.0258 1.0000 0.0506 -7.500 -0.4402 0.08127 0.07810 -0.0234 1.0000 0.0510 -7.250 -0.4364 0.07776 0.07462 -0.0222 1.0000 0.0514 -7.000 -0.4327 0.07445 0.07131 -0.0218 1.0000 0.0517 -6.750 -0.4285 0.07112 0.06796 -0.0219 1.0000 0.0520 -6.500 -0.4237 0.06744 0.06426 -0.0226 1.0000 0.0521 -6.250 -0.4203 0.05987 0.05658 -0.0257 1.0000 0.0394 -6.000 -0.4179 0.05100 0.04754 -0.0270 1.0000 0.0303 -5.750 -0.4129 0.04683 0.04324 -0.0259 1.0000 0.0290 -5.500 -0.4054 0.04176 0.03796 -0.0250 1.0000 0.0280 -5.250 -0.4010 0.03437 0.03014 -0.0232 1.0000 0.0264 -5.000 -0.4093 0.02100 0.01519 -0.0177 1.0000 0.0243 -4.750 -0.3919 0.01890 0.01260 -0.0153 1.0000 0.0249 -4.500 -0.3725 0.01767 0.01106 -0.0135 1.0000 0.0281 -4.250 -0.3517 0.01653 0.00958 -0.0116 1.0000 0.0297 -4.000 -0.3302 0.01581 0.00862 -0.0099 1.0000 0.0319 -3.750 -0.3084 0.01567 0.00860 -0.0083 1.0000 0.0392 -3.500 -0.2861 0.01689 0.00984 -0.0068 1.0000 0.0620 -3.250 -0.2660 0.01942 0.01255 -0.0054 1.0000 0.0761 -3.000 -0.2459 0.02067 0.01370 -0.0041 0.9999 0.0875 -2.750 -0.2051 0.02161 0.01496 -0.0082 0.9942 0.1062 -1.750 -0.0499 0.02024 0.01304 -0.0185 0.9698 0.1523 -1.500 -0.0141 0.01922 0.01169 -0.0195 0.9621 0.1397 -1.250 0.0294 0.01787 0.01031 -0.0226 0.9572 0.1352 -1.000 0.0667 0.01697 0.00924 -0.0241 0.9492 0.1305 -0.750 0.1125 0.01626 0.00837 -0.0274 0.9436 0.1282 -0.500 0.1500 0.01562 0.00770 -0.0290 0.9345 0.1279 -0.250 0.1986 0.01490 0.00699 -0.0330 0.9293 0.1279 0.000 0.2367 0.01426 0.00637 -0.0346 0.9189 0.1283 0.250 0.2797 0.01356 0.00576 -0.0373 0.9097 0.1295 0.500 0.3274 0.01287 0.00518 -0.0410 0.9000 0.1327 0.750 0.3645 0.01231 0.00470 -0.0424 0.8818 0.1338 1.000 0.4062 0.01175 0.00420 -0.0447 0.8584 0.1348 1.250 0.4418 0.01135 0.00382 -0.0457 0.8223 0.1367 1.500 0.4771 0.01109 0.00347 -0.0466 0.7676 0.1388 1.750 0.5061 0.01104 0.00321 -0.0463 0.6971 0.1420 2.000 0.5289 0.01128 0.00316 -0.0449 0.6397 0.1460 2.250 0.5508 0.01156 0.00323 -0.0435 0.5969 0.1505 2.500 0.5718 0.01183 0.00333 -0.0419 0.5538 0.1566 2.750 0.5905 0.01211 0.00340 -0.0399 0.4957 0.1698 3.000 0.7390 0.01153 0.00409 -0.0684 0.3814 1.0000 3.250 0.7605 0.01194 0.00426 -0.0672 0.3317 1.0000 3.500 0.7812 0.01248 0.00445 -0.0659 0.2632 1.0000 3.750 0.8023 0.01308 0.00475 -0.0647 0.2043 1.0000 4.000 0.8239 0.01371 0.00514 -0.0635 0.1565 1.0000 4.250 0.8390 0.01522 0.00594 -0.0614 0.0382 1.0000 4.500 0.8601 0.01599 0.00667 -0.0600 0.0261 1.0000 4.750 0.8827 0.01657 0.00739 -0.0588 0.0245 1.0000 5.000 0.9050 0.01720 0.00818 -0.0575 0.0238 1.0000 5.250 0.9266 0.01792 0.00912 -0.0561 0.0236 1.0000 5.500 0.9467 0.01878 0.01016 -0.0544 0.0236 1.0000 5.750 0.9657 0.01974 0.01126 -0.0525 0.0241 1.0000 6.000 0.9831 0.02091 0.01255 -0.0503 0.0248 1.0000 6.250 1.0002 0.02234 0.01406 -0.0480 0.0259 1.0000 6.500 1.0185 0.02393 0.01573 -0.0460 0.0260 1.0000 6.750 1.0380 0.02557 0.01745 -0.0444 0.0251 1.0000 7.000 1.0599 0.02788 0.01995 -0.0428 0.0269 1.0000 7.250 1.0834 0.03205 0.02426 -0.0417 0.0295 1.0000 7.500 1.1059 0.03393 0.02684 -0.0388 0.0360 1.0000 7.750 1.1243 0.03848 0.03212 -0.0355 0.0472 1.0000 14.000 0.8998 0.18048 0.17701 -0.0572 0.0506 1.0000 14.250 0.8953 0.18577 0.18226 -0.0624 0.0498 1.0000