XFOIL Version 6.94 Calculated polar for: GOE 122 (MVA H.2) 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 ------- -------- --------- --------- -------- ------- ------- 0.000 0.3842 0.01100 0.00342 -0.0662 0.6982 0.0364 0.500 0.4379 0.01092 0.00314 -0.0658 0.6550 0.0512 1.000 0.4924 0.01063 0.00269 -0.0652 0.6141 0.0499 1.500 0.5473 0.01061 0.00243 -0.0648 0.5707 0.0502 2.000 0.6019 0.01072 0.00232 -0.0644 0.5256 0.0576 2.500 0.6533 0.00899 0.00243 -0.0637 0.4799 1.0000 3.000 0.7073 0.00945 0.00265 -0.0634 0.4499 1.0000 3.500 0.7614 0.00988 0.00299 -0.0632 0.4249 1.0000 4.000 0.8151 0.01033 0.00333 -0.0629 0.3991 1.0000 4.500 0.8679 0.01079 0.00366 -0.0625 0.3480 1.0000 5.000 0.9185 0.01162 0.00407 -0.0621 0.2542 1.0000 5.500 0.9643 0.01334 0.00511 -0.0613 0.1399 1.0000 6.000 1.0074 0.01546 0.00658 -0.0601 0.0315 1.0000 6.500 1.0574 0.01632 0.00746 -0.0595 0.0164 1.0000 7.000 1.1058 0.01743 0.00869 -0.0585 0.0041 1.0000 7.500 1.1541 0.01849 0.00999 -0.0574 0.0038 1.0000 8.000 1.1992 0.01990 0.01166 -0.0558 0.0038 1.0000 8.500 1.2393 0.02181 0.01390 -0.0537 0.0040 1.0000 9.000 1.2703 0.02447 0.01697 -0.0505 0.0043 1.0000 9.500 1.2874 0.02809 0.02109 -0.0459 0.0047 1.0000 10.000 1.2882 0.03255 0.02594 -0.0397 0.0050 1.0000 10.500 1.2910 0.03753 0.03127 -0.0351 0.0053 1.0000 11.000 1.3020 0.04182 0.03587 -0.0323 0.0059 1.0000 11.500 1.3033 0.04882 0.04341 -0.0286 0.0070 1.0000 12.000 1.2830 0.05928 0.05453 -0.0261 0.0082 1.0000 12.500 1.2534 0.06993 0.06567 -0.0277 0.0087 1.0000 13.000 1.2201 0.08189 0.07804 -0.0324 0.0090 1.0000 13.500 1.1858 0.09525 0.09175 -0.0394 0.0091 1.0000 14.000 1.1533 0.10996 0.10677 -0.0482 0.0089 1.0000 14.500 1.1216 0.12651 0.12359 -0.0588 0.0087 1.0000 15.000 1.0933 0.14461 0.14190 -0.0706 0.0083 1.0000 15.500 1.0674 0.16519 0.16260 -0.0832 0.0078 1.0000 16.000 1.0490 0.18655 0.18392 -0.0940 0.0078 1.0000 16.500 1.0481 0.20111 0.19842 -0.1016 0.0088 1.0000