XFOIL Version 6.94 Calculated polar for: GOE 133 (MVA H.11) AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.3148 0.00924 0.00206 -0.0623 0.6615 0.0902 0.500 0.3683 0.00927 0.00190 -0.0615 0.5935 0.1065 1.000 0.4183 0.00795 0.00195 -0.0611 0.5456 0.6205 1.500 0.4775 0.00735 0.00206 -0.0609 0.5140 1.0000 2.000 0.5320 0.00766 0.00217 -0.0604 0.4890 1.0000 2.500 0.5868 0.00795 0.00232 -0.0599 0.4672 1.0000 3.000 0.6418 0.00822 0.00250 -0.0595 0.4475 1.0000 3.500 0.6965 0.00851 0.00270 -0.0591 0.4243 1.0000 4.000 0.7511 0.00882 0.00294 -0.0586 0.3990 1.0000 4.500 0.8052 0.00917 0.00319 -0.0581 0.3637 1.0000 5.000 0.8576 0.00973 0.00352 -0.0574 0.3016 1.0000 5.500 0.9072 0.01069 0.00409 -0.0565 0.2339 1.0000 6.000 0.9584 0.01145 0.00472 -0.0558 0.2090 1.0000 6.500 1.0099 0.01211 0.00535 -0.0550 0.1892 1.0000 7.000 1.0613 0.01275 0.00597 -0.0543 0.1724 1.0000 7.500 1.1123 0.01341 0.00660 -0.0535 0.1431 1.0000 8.000 1.1537 0.01526 0.00796 -0.0517 0.0677 1.0000 8.500 1.1966 0.01681 0.00953 -0.0498 0.0540 1.0000 9.000 1.2384 0.01831 0.01107 -0.0479 0.0464 1.0000 9.500 1.2792 0.01976 0.01261 -0.0458 0.0412 1.0000 10.000 1.3157 0.02145 0.01439 -0.0432 0.0373 1.0000 10.500 1.3507 0.02308 0.01614 -0.0405 0.0345 1.0000 11.000 1.3781 0.02512 0.01834 -0.0369 0.0324 1.0000 11.500 1.3992 0.02709 0.02041 -0.0326 0.0304 1.0000 12.000 1.4139 0.02973 0.02323 -0.0283 0.0289 1.0000 12.500 1.4280 0.03263 0.02632 -0.0249 0.0274 1.0000 13.000 1.4368 0.03635 0.03013 -0.0222 0.0263 1.0000 13.500 1.4437 0.04070 0.03477 -0.0203 0.0253 1.0000 14.000 1.4488 0.04561 0.03994 -0.0195 0.0242 1.0000 14.500 1.4519 0.05102 0.04552 -0.0195 0.0234 1.0000 15.000 1.4508 0.05700 0.05163 -0.0191 0.0225 1.0000 15.500 1.4404 0.06497 0.06000 -0.0216 0.0218 1.0000 16.000 1.4291 0.07339 0.06871 -0.0246 0.0210 1.0000 16.500 1.4176 0.08221 0.07774 -0.0281 0.0203 1.0000 17.000 1.4061 0.09100 0.08671 -0.0313 0.0199 1.0000 17.500 1.3937 0.09967 0.09553 -0.0336 0.0194 1.0000 18.000 1.3587 0.11452 0.11079 -0.0425 0.0189 1.0000 18.500 1.3312 0.12855 0.12514 -0.0508 0.0182 1.0000 19.000 1.2905 0.14550 0.14247 -0.0603 0.0185 1.0000