XFOIL Version 6.96 Calculated polar for: WORTMANN FX 05-188 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -16.500 -0.5840 0.10069 0.09860 -0.0776 0.9546 0.0177 -16.250 -0.5973 0.08979 0.08752 -0.0884 0.9422 0.0179 -16.000 -0.6039 0.08054 0.07805 -0.0973 0.9262 0.0164 -15.750 -0.6244 0.06995 0.06711 -0.1079 0.9049 0.0164 -15.500 -0.6424 0.06110 0.05792 -0.1166 0.8888 0.0161 -15.250 -0.6500 0.05506 0.05157 -0.1227 0.8750 0.0162 -15.000 -0.6599 0.04990 0.04612 -0.1265 0.8637 0.0161 -14.750 -0.6657 0.04602 0.04201 -0.1287 0.8552 0.0161 -14.500 -0.6704 0.04253 0.03827 -0.1302 0.8475 0.0161 -14.250 -0.6705 0.03977 0.03531 -0.1310 0.8409 0.0160 -14.000 -0.6686 0.03728 0.03264 -0.1312 0.8349 0.0160 -13.750 -0.6615 0.03543 0.03061 -0.1316 0.8288 0.0162 -13.500 -0.6542 0.03376 0.02881 -0.1314 0.8236 0.0164 -13.250 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0.00430 -0.1186 0.4843 0.7164 6.250 1.1579 0.00947 0.00457 -0.1170 0.4556 0.7175 6.750 1.1791 0.01048 0.00528 -0.1111 0.3897 0.7211 7.000 1.1860 0.01103 0.00571 -0.1074 0.3602 0.7227 7.250 1.1882 0.01181 0.00631 -0.1031 0.3233 0.7242 7.500 1.1826 0.01290 0.00717 -0.0978 0.2757 0.7258 7.750 1.1758 0.01416 0.00820 -0.0926 0.2315 0.7274 8.000 1.1704 0.01552 0.00937 -0.0882 0.1922 0.7291 8.250 1.1692 0.01687 0.01055 -0.0846 0.1598 0.7308 8.500 1.1698 0.01825 0.01178 -0.0815 0.1301 0.7323 8.750 1.1692 0.01979 0.01315 -0.0785 0.0996 0.7337 9.000 1.1759 0.02101 0.01429 -0.0764 0.0849 0.7348 9.250 1.1864 0.02201 0.01528 -0.0749 0.0761 0.7369 9.500 1.1976 0.02300 0.01627 -0.0735 0.0681 0.7386 9.750 1.2071 0.02413 0.01738 -0.0719 0.0586 0.7400 10.000 1.2124 0.02558 0.01873 -0.0701 0.0434 0.7414 10.250 1.2181 0.02703 0.02013 -0.0683 0.0334 0.7427 10.500 1.2285 0.02823 0.02133 -0.0671 0.0302 0.7439 10.750 1.2391 0.02945 0.02257 -0.0660 0.0286 0.7450 11.000 1.2487 0.03076 0.02391 -0.0648 0.0270 0.7461 11.250 1.2598 0.03200 0.02519 -0.0638 0.0262 0.7473 11.500 1.2719 0.03320 0.02644 -0.0630 0.0256 0.7485 11.750 1.2830 0.03451 0.02779 -0.0621 0.0250 0.7498 12.000 1.2938 0.03587 0.02919 -0.0613 0.0243 0.7508 12.250 1.3033 0.03737 0.03072 -0.0604 0.0237 0.7518 12.500 1.3117 0.03898 0.03238 -0.0596 0.0230 0.7533 12.750 1.3167 0.04096 0.03443 -0.0586 0.0222 0.7546 13.000 1.3260 0.04258 0.03612 -0.0579 0.0219 0.7560 13.250 1.3362 0.04416 0.03776 -0.0573 0.0217 0.7574 13.500 1.3455 0.04585 0.03951 -0.0568 0.0214 0.7588 13.750 1.3551 0.04755 0.04127 -0.0563 0.0211 0.7602 14.000 1.3628 0.04948 0.04327 -0.0558 0.0207 0.7617 14.250 1.3711 0.05138 0.04523 -0.0554 0.0204 0.7632 14.500 1.3784 0.05344 0.04735 -0.0551 0.0201 0.7648 14.750 1.3856 0.05553 0.04949 -0.0548 0.0197 0.7664 15.000 1.3911 0.05787 0.05188 -0.0545 0.0194 0.7679 15.250 1.3961 0.06029 0.05437 -0.0544 0.0191 0.7698 15.500 1.3970 0.06325 0.05740 -0.0542 0.0186 0.7716 15.750 1.3944 0.06669 0.06094 -0.0541 0.0183 0.7731 16.000 1.3982 0.06942 0.06375 -0.0541 0.0181 0.7750 16.250 1.4057 0.07177 0.06617 -0.0544 0.0179 0.7770 16.500 1.4125 0.07422 0.06869 -0.0546 0.0178 0.7792 16.750 1.4181 0.07688 0.07143 -0.0550 0.0176 0.7813 17.000 1.4239 0.07952 0.07413 -0.0553 0.0174 0.7834 17.250 1.4283 0.08239 0.07709 -0.0558 0.0171 0.7856 17.500 1.4331 0.08521 0.08000 -0.0564 0.0169 0.7879 17.750 1.4371 0.08822 0.08309 -0.0570 0.0167 0.7904 18.000 1.4418 0.09113 0.08608 -0.0577 0.0164 0.7930 18.250 1.4453 0.09427 0.08928 -0.0585 0.0162 0.7957 18.500 1.4488 0.09740 0.09247 -0.0594 0.0160 0.7984 18.750 1.4512 0.10073 0.09587 -0.0604 0.0158 0.8009 19.000 1.4532 0.10412 0.09934 -0.0615 0.0156 0.8037 19.250 1.4544 0.10766 0.10296 -0.0627 0.0154 0.8066