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 = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.2556 0.01000 0.00251 -0.0427 0.6269 0.0518 0.500 0.3110 0.00984 0.00223 -0.0423 0.5814 0.0532 1.000 0.3666 0.00967 0.00194 -0.0420 0.5363 0.0633 1.500 0.4223 0.00974 0.00189 -0.0417 0.4886 0.0717 2.000 0.4814 0.00759 0.00205 -0.0427 0.4512 1.0000 2.500 0.5359 0.00795 0.00218 -0.0422 0.4253 1.0000 3.000 0.5906 0.00829 0.00237 -0.0418 0.4030 1.0000 3.500 0.6454 0.00861 0.00260 -0.0415 0.3851 1.0000 4.500 0.7551 0.00925 0.00309 -0.0408 0.3404 1.0000 5.000 0.8099 0.00958 0.00340 -0.0405 0.3199 1.0000 5.500 0.8641 0.00997 0.00374 -0.0401 0.2877 1.0000 6.000 0.9159 0.01077 0.00420 -0.0396 0.2141 1.0000 6.500 0.9644 0.01210 0.00514 -0.0389 0.1468 1.0000 7.000 1.0085 0.01416 0.00661 -0.0377 0.0440 1.0000 7.500 1.0564 0.01546 0.00796 -0.0366 0.0354 1.0000 8.000 1.1021 0.01697 0.00957 -0.0353 0.0329 1.0000 8.500 1.1458 0.01856 0.01126 -0.0338 0.0305 1.0000 9.000 1.1845 0.02052 0.01333 -0.0318 0.0290 1.0000 9.500 1.2229 0.02231 0.01523 -0.0298 0.0268 1.0000 10.000 1.2515 0.02509 0.01802 -0.0269 0.0256 1.0000 10.500 1.2848 0.02718 0.02032 -0.0243 0.0245 1.0000 11.000 1.3135 0.02952 0.02280 -0.0214 0.0232 1.0000 11.500 1.3399 0.03250 0.02580 -0.0184 0.0223 1.0000 12.000 1.3636 0.03588 0.02942 -0.0154 0.0217 1.0000 12.500 1.3769 0.03938 0.03326 -0.0121 0.0210 1.0000 13.000 1.3859 0.04352 0.03773 -0.0095 0.0204 1.0000 13.500 1.3952 0.04756 0.04192 -0.0081 0.0194 1.0000 14.000 1.4022 0.05319 0.04772 -0.0067 0.0185 1.0000 14.500 1.3803 0.06061 0.05562 -0.0077 0.0183 1.0000 15.000 1.3517 0.07016 0.06563 -0.0114 0.0178 1.0000 15.500 1.3227 0.08099 0.07686 -0.0162 0.0176 1.0000 16.000 1.2900 0.09346 0.08969 -0.0228 0.0173 1.0000 16.500 1.2503 0.10866 0.10526 -0.0316 0.0174 1.0000 17.000 1.2063 0.12631 0.12326 -0.0425 0.0174 1.0000 18.000 1.1021 0.17242 0.16993 -0.0714 0.0184 1.0000