XFOIL Version 6.96 Calculated polar for: GIII BL167 AIRFOIL 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 ------ -------- --------- --------- -------- -------- -------- -9.000 -0.4629 0.08640 0.08316 -0.0135 1.0000 0.0414 -8.750 -0.4654 0.08207 0.07886 -0.0149 1.0000 0.0424 -8.500 -0.4696 0.07752 0.07433 -0.0169 1.0000 0.0432 -8.250 -0.4760 0.07273 0.06958 -0.0194 1.0000 0.0440 -8.000 -0.5823 0.07671 0.07335 -0.0217 1.0000 0.0394 -7.750 -0.5777 0.07309 0.06970 -0.0226 1.0000 0.0406 -7.500 -0.5730 0.06911 0.06567 -0.0241 1.0000 0.0418 -7.250 -0.5670 0.06506 0.06153 -0.0255 1.0000 0.0433 -7.000 -0.5590 0.06095 0.05728 -0.0267 1.0000 0.0455 -6.750 -0.5452 0.05891 0.05467 -0.0274 1.0000 0.0492 -5.250 -0.4866 0.03699 0.03175 -0.0194 1.0000 0.0646 -5.000 -0.4755 0.03461 0.02937 -0.0174 1.0000 0.0668 -4.750 -0.4662 0.03342 0.02776 -0.0144 1.0000 0.0776 -4.500 -0.4467 0.02665 0.02038 -0.0109 1.0000 0.0451 -4.250 -0.4262 0.02281 0.01568 -0.0079 1.0000 0.0376 -4.000 -0.4059 0.02059 0.01314 -0.0064 1.0000 0.0378 -3.750 -0.3821 0.01844 0.01080 -0.0058 0.9996 0.0395 -3.500 -0.3440 0.01749 0.00975 -0.0082 0.9958 0.0452 -3.250 -0.3062 0.01624 0.00827 -0.0099 0.9923 0.0488 -3.000 -0.2709 0.01478 0.00684 -0.0117 0.9880 0.0570 -2.750 -0.2329 0.01388 0.00596 -0.0140 0.9839 0.0723 -2.500 -0.1978 0.01309 0.00523 -0.0160 0.9784 0.0937 -2.250 -0.1612 0.01240 0.00465 -0.0182 0.9729 0.1296 -2.000 -0.1391 0.00954 0.00434 -0.0181 0.9692 0.7248 -1.750 -0.1158 0.00947 0.00459 -0.0156 0.9616 0.8515 -1.500 -0.0831 0.00962 0.00473 -0.0153 0.9571 0.9004 -1.250 -0.0454 0.00984 0.00487 -0.0162 0.9535 0.9328 -1.000 -0.0025 0.01003 0.00496 -0.0186 0.9496 0.9552 -0.750 0.0805 0.01024 0.00506 -0.0289 0.9579 0.9790 -0.500 0.1386 0.01013 0.00486 -0.0354 0.9566 0.9883 -0.250 0.1800 0.01000 0.00469 -0.0386 0.9482 0.9923 0.000 0.2207 0.00982 0.00448 -0.0415 0.9377 0.9957 0.250 0.2581 0.00963 0.00425 -0.0435 0.9243 1.0000 0.500 0.2792 0.00954 0.00413 -0.0422 0.9083 1.0000 0.750 0.3004 0.00948 0.00405 -0.0409 0.8945 1.0000 1.000 0.3220 0.00945 0.00400 -0.0398 0.8821 1.0000 1.250 0.3440 0.00942 0.00396 -0.0387 0.8703 1.0000 1.500 0.3663 0.00941 0.00395 -0.0377 0.8582 1.0000 1.750 0.3887 0.00941 0.00397 -0.0368 0.8460 1.0000 2.000 0.4111 0.00943 0.00400 -0.0358 0.8339 1.0000 2.250 0.4337 0.00946 0.00407 -0.0347 0.8217 1.0000 2.500 0.4563 0.00949 0.00413 -0.0336 0.8092 1.0000 2.750 0.4783 0.00949 0.00414 -0.0322 0.7933 1.0000 3.000 0.4986 0.00939 0.00400 -0.0301 0.7676 1.0000 3.250 0.5196 0.00932 0.00391 -0.0282 0.7394 1.0000 3.500 0.5416 0.00931 0.00392 -0.0267 0.7111 1.0000 3.750 0.5633 0.00933 0.00389 -0.0250 0.6741 1.0000 4.000 0.5843 0.00945 0.00388 -0.0233 0.6211 1.0000 4.250 0.6032 0.00980 0.00393 -0.0213 0.5274 1.0000 4.500 0.6155 0.01093 0.00423 -0.0185 0.3635 1.0000 4.750 0.6218 0.01321 0.00515 -0.0157 0.1148 1.0000 5.000 0.6379 0.01465 0.00629 -0.0138 0.0697 1.0000 5.250 0.6564 0.01577 0.00739 -0.0123 0.0576 1.0000 5.500 0.6774 0.01658 0.00822 -0.0111 0.0493 1.0000 5.750 0.6961 0.01811 0.00976 -0.0096 0.0452 1.0000 6.000 0.7186 0.01928 0.01101 -0.0084 0.0420 1.0000 6.250 0.7411 0.02030 0.01209 -0.0076 0.0374 1.0000 6.500 0.7367 0.01118 0.00342 -0.0042 0.0350 1.0000 6.750 0.7589 0.01330 0.00582 -0.0031 0.0340 1.0000 7.000 0.7797 0.01526 0.00812 -0.0018 0.0332 1.0000 7.250 0.7990 0.01693 0.01014 -0.0003 0.0309 1.0000 7.500 0.8148 0.02000 0.01364 0.0015 0.0307 1.0000 7.750 0.8258 0.02410 0.01819 0.0037 0.0319 1.0000 11.250 0.6274 0.10897 0.10569 -0.0027 0.0570 1.0000 11.500 0.6148 0.11503 0.11171 -0.0062 0.0551 1.0000