XFOIL Version 6.94 Calculated polar for: WORTMANN FX 08-S-176 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 1.500 0.6257 0.02102 0.01518 -0.0711 0.5635 0.9487 2.000 0.6748 0.02143 0.01547 -0.0708 0.5607 0.9560 2.500 0.7270 0.02222 0.01640 -0.0723 0.5553 0.9619 3.000 0.7603 0.02281 0.01704 -0.0696 0.5498 0.9700 3.500 0.8198 0.02272 0.01690 -0.0714 0.5463 0.9736 4.000 0.8905 0.02232 0.01646 -0.0751 0.5434 0.9763 4.500 0.9657 0.02191 0.01598 -0.0795 0.5410 0.9799 5.000 1.0458 0.01859 0.01249 -0.0829 0.5261 0.9834 5.500 1.1135 0.01615 0.00988 -0.0845 0.5105 0.9877 6.500 1.2271 0.01472 0.00838 -0.0867 0.4635 0.9959 7.000 1.2581 0.01503 0.00854 -0.0834 0.4251 1.0000 7.500 1.2135 0.01564 0.00908 -0.0657 0.4013 1.0000 8.000 1.1612 0.01673 0.01019 -0.0486 0.3837 1.0000 8.500 1.1267 0.02097 0.01434 -0.0412 0.3487 1.0000 9.000 1.1090 0.02627 0.01960 -0.0382 0.3227 1.0000 9.500 1.0950 0.03136 0.02464 -0.0354 0.2929 1.0000 10.500 0.9906 0.04848 0.04085 -0.0272 0.1242 1.0000 11.000 0.9526 0.05733 0.04921 -0.0252 0.0179 1.0000 11.500 0.9645 0.06184 0.05377 -0.0252 0.0051 1.0000 12.000 0.9836 0.06570 0.05777 -0.0255 0.0048 1.0000 12.500 0.9990 0.07012 0.06236 -0.0259 0.0047 1.0000 13.000 1.0140 0.07467 0.06709 -0.0265 0.0047 1.0000 13.500 1.0253 0.07986 0.07247 -0.0272 0.0047 1.0000 14.000 1.0349 0.08535 0.07816 -0.0281 0.0048 1.0000 14.500 1.0410 0.09155 0.08455 -0.0293 0.0049 1.0000 15.000 1.0444 0.09825 0.09143 -0.0308 0.0050 1.0000 15.500 1.0476 0.10509 0.09842 -0.0325 0.0052 1.0000 16.000 1.0522 0.11171 0.10514 -0.0343 0.0053 1.0000 16.500 1.0649 0.11675 0.11021 -0.0356 0.0055 1.0000 17.000 1.0854 0.12102 0.11464 -0.0371 0.0058 1.0000 17.500 1.1176 0.12263 0.11633 -0.0370 0.0066 1.0000 18.000 1.1689 0.12089 0.11459 -0.0353 0.0075 1.0000