XFOIL Version 6.96 Calculated polar for: WORTMANN FX 77-W-153 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 = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -12.250 -0.9075 0.03948 0.03547 -0.0397 0.4740 0.0092 -12.000 -0.9177 0.03601 0.03175 -0.0378 0.4696 0.0092 -11.750 -0.9149 0.03339 0.02889 -0.0363 0.4631 0.0092 -11.500 -0.9071 0.03118 0.02644 -0.0350 0.4568 0.0092 -11.250 -0.8947 0.02940 0.02446 -0.0337 0.4495 0.0092 -11.000 -0.8787 0.02801 0.02289 -0.0327 0.4417 0.0093 -10.750 -0.8608 0.02679 0.02150 -0.0317 0.4351 0.0093 -10.500 -0.8412 0.02575 0.02031 -0.0308 0.4285 0.0094 -10.250 -0.8201 0.02488 0.01930 -0.0301 0.4234 0.0094 -10.000 -0.7986 0.02402 0.01831 -0.0293 0.4191 0.0095 -9.750 -0.7762 0.02325 0.01742 -0.0286 0.4149 0.0095 -9.500 -0.7539 0.02243 0.01646 -0.0279 0.4108 0.0096 -9.250 -0.7328 0.02136 0.01524 -0.0270 0.4079 0.0096 -9.000 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0.3287 0.9999 6.250 1.0620 0.00894 0.00373 -0.0389 0.3261 1.0000 6.500 1.0887 0.00898 0.00377 -0.0388 0.3205 1.0000 6.750 1.1151 0.00904 0.00380 -0.0386 0.3121 1.0000 7.000 1.1406 0.00920 0.00386 -0.0384 0.2959 1.0000 7.250 1.1652 0.00947 0.00406 -0.0380 0.2839 1.0000 7.500 1.1898 0.00973 0.00428 -0.0376 0.2754 1.0000 7.750 1.2146 0.00994 0.00448 -0.0373 0.2702 1.0000 8.000 1.2392 0.01016 0.00469 -0.0369 0.2615 1.0000 8.250 1.2625 0.01050 0.00497 -0.0364 0.2499 1.0000 8.500 1.2848 0.01091 0.00531 -0.0357 0.2378 1.0000 8.750 1.3070 0.01129 0.00566 -0.0351 0.2292 1.0000 9.000 1.3301 0.01157 0.00594 -0.0345 0.2226 1.0000 9.250 1.3508 0.01203 0.00634 -0.0336 0.2120 1.0000 9.500 1.3708 0.01251 0.00677 -0.0326 0.1970 1.0000 9.750 1.3884 0.01314 0.00731 -0.0313 0.1819 1.0000 10.000 1.4066 0.01367 0.00779 -0.0301 0.1718 1.0000 10.250 1.4261 0.01408 0.00820 -0.0290 0.1648 1.0000 10.500 1.4391 0.01486 0.00890 -0.0271 0.1454 1.0000 10.750 1.4467 0.01588 0.00979 -0.0245 0.1267 1.0000 11.000 1.4462 0.01720 0.01100 -0.0208 0.1005 1.0000 11.250 1.4495 0.01821 0.01197 -0.0177 0.0901 1.0000 11.500 1.4516 0.01909 0.01285 -0.0144 0.0821 1.0000 11.750 1.4486 0.02047 0.01424 -0.0112 0.0716 1.0000 12.000 1.4391 0.02299 0.01675 -0.0094 0.0599 1.0000 12.250 1.4360 0.02596 0.01979 -0.0098 0.0531 1.0000 12.500 1.4300 0.02997 0.02386 -0.0113 0.0477 1.0000 12.750 1.4245 0.03396 0.02794 -0.0126 0.0433 1.0000 13.000 1.4027 0.03968 0.03374 -0.0138 0.0360 1.0000 13.250 1.3807 0.04514 0.03927 -0.0146 0.0319 1.0000 13.500 1.3668 0.04954 0.04375 -0.0151 0.0303 1.0000 13.750 1.3521 0.05403 0.04831 -0.0156 0.0284 1.0000 14.000 1.3378 0.05844 0.05279 -0.0161 0.0266 1.0000 14.250 1.3305 0.06215 0.05657 -0.0165 0.0251 1.0000 14.500 1.3215 0.06619 0.06066 -0.0171 0.0221 1.0000 14.750 1.3073 0.07098 0.06549 -0.0180 0.0167 1.0000 15.000 1.2982 0.07517 0.06970 -0.0187 0.0144 1.0000 15.250 1.2941 0.07880 0.07338 -0.0194 0.0137 1.0000 15.500 1.2918 0.08221 0.07684 -0.0200 0.0131 1.0000 15.750 1.2882 0.08584 0.08053 -0.0208 0.0127 1.0000