XFOIL Version 6.94 Calculated polar for: GD BAP 00 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1369 0.05181 0.04285 -0.0503 0.9998 0.2831 -2.750 -0.0994 0.04909 0.03998 -0.0538 0.9998 0.2728 -2.500 -0.0572 0.04690 0.03748 -0.0581 0.9998 0.2615 -2.250 -0.0146 0.04548 0.03571 -0.0617 0.9998 0.2524 -2.000 0.0236 0.04481 0.03497 -0.0641 0.9998 0.2477 -1.750 0.1304 0.04018 0.03017 -0.0800 0.8796 0.2459 -1.500 0.3079 0.03468 0.02256 -0.1054 0.6765 0.2652 -1.250 0.3716 0.03479 0.02191 -0.1119 0.6550 0.2765 -1.000 0.4293 0.03528 0.02193 -0.1174 0.6434 0.2912 -0.750 0.4712 0.03519 0.02212 -0.1202 0.6369 0.3199 -0.500 0.5102 0.03471 0.02286 -0.1224 0.6326 0.4401 -0.250 0.5226 0.03272 0.02170 -0.1160 0.6312 1.0002 0.000 0.5602 0.03357 0.02192 -0.1174 0.6299 1.0002 0.250 0.5953 0.03456 0.02261 -0.1191 0.6276 1.0002 0.500 0.6296 0.03590 0.02376 -0.1210 0.6234 1.0002 0.750 0.6633 0.03783 0.02550 -0.1230 0.6179 1.0002 1.000 0.6915 0.03920 0.02691 -0.1240 0.6133 1.0002 1.250 0.7205 0.04066 0.02835 -0.1252 0.6109 1.0002 1.500 0.7489 0.04217 0.02986 -0.1264 0.6095 1.0002 1.750 0.7760 0.04382 0.03153 -0.1275 0.6082 1.0002 2.000 0.8021 0.04565 0.03340 -0.1285 0.6068 1.0002 2.250 0.8270 0.04770 0.03549 -0.1294 0.6055 1.0002 2.500 0.8485 0.04967 0.03755 -0.1299 0.6043 1.0002 2.750 0.8647 0.05184 0.03985 -0.1298 0.6016 1.0002 3.000 0.8789 0.05460 0.04270 -0.1293 0.5966 1.0002 3.250 0.8888 0.05805 0.04626 -0.1282 0.5891 1.0002 3.500 0.8971 0.06198 0.05022 -0.1267 0.5791 1.0002 3.750 0.8966 0.06615 0.05445 -0.1243 0.5666 1.0002 4.000 0.8982 0.07050 0.05884 -0.1222 0.5549 1.0002 4.250 0.8969 0.07453 0.06292 -0.1203 0.5477 1.0002 4.500 0.8978 0.07849 0.06693 -0.1190 0.5437 1.0002 4.750 0.8879 0.08243 0.07095 -0.1170 0.5407 1.0002 5.000 0.8911 0.08662 0.07515 -0.1163 0.5379 1.0002