XFOIL Version 6.96 Calculated polar for: GOE 590 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 ------ -------- --------- --------- -------- -------- -------- -8.500 -0.4078 0.09766 0.09612 -0.0098 1.0000 0.0034 -8.250 -0.4035 0.09460 0.09307 -0.0110 1.0000 0.0034 -8.000 -0.3992 0.09159 0.09008 -0.0122 0.9996 0.0034 -7.750 -0.3485 0.08449 0.08285 -0.0279 0.9233 0.0036 -5.750 -0.2778 0.05912 0.05646 -0.0430 0.6879 0.0029 -5.500 -0.2602 0.05547 0.05273 -0.0451 0.6782 0.0027 -5.250 -0.2409 0.05158 0.04873 -0.0471 0.6689 0.0026 -5.000 -0.2202 0.04773 0.04476 -0.0487 0.6587 0.0025 -4.750 -0.1984 0.04358 0.04046 -0.0501 0.6471 0.0025 -4.500 -0.1755 0.03913 0.03583 -0.0510 0.6343 0.0024 -4.000 -0.1431 0.01358 0.00849 -0.0467 0.6226 0.0029 -3.750 -0.1188 0.01201 0.00649 -0.0460 0.6036 0.0031 -3.500 -0.0935 0.01112 0.00529 -0.0455 0.5846 0.0032 -3.250 -0.0682 0.01027 0.00420 -0.0450 0.5691 0.0038 -3.000 -0.0422 0.00990 0.00369 -0.0446 0.5560 0.0042 -2.750 -0.0163 0.00942 0.00304 -0.0442 0.5450 0.0047 -2.250 0.0356 0.00869 0.00205 -0.0433 0.5274 0.0070 -1.750 0.0883 0.00832 0.00160 -0.0427 0.5147 0.0156 -1.500 0.1152 0.00829 0.00156 -0.0425 0.5091 0.0213 -1.250 0.1425 0.00833 0.00157 -0.0425 0.5039 0.0260 -1.000 0.1699 0.00837 0.00158 -0.0424 0.4988 0.0278 -0.750 0.1969 0.00836 0.00153 -0.0423 0.4938 0.0290 -0.500 0.2231 0.00818 0.00128 -0.0420 0.4898 0.0301 -0.250 0.2497 0.00806 0.00113 -0.0418 0.4860 0.0307 0.000 0.2764 0.00798 0.00101 -0.0416 0.4819 0.0311 0.250 0.3031 0.00792 0.00092 -0.0414 0.4779 0.0319 0.500 0.3300 0.00787 0.00085 -0.0412 0.4742 0.0325 0.750 0.3569 0.00783 0.00079 -0.0411 0.4704 0.0326 1.000 0.3838 0.00780 0.00075 -0.0410 0.4666 0.0329 1.250 0.4106 0.00780 0.00073 -0.0408 0.4632 0.0335 1.500 0.4377 0.00778 0.00072 -0.0407 0.4601 0.0349 2.000 0.4908 0.00771 0.00078 -0.0404 0.4465 0.0887 2.250 0.5173 0.00766 0.00084 -0.0402 0.4412 0.1336 2.500 0.5437 0.00761 0.00090 -0.0401 0.4368 0.1862 3.000 0.6133 0.00599 0.00121 -0.0440 0.4222 0.9804 3.250 0.6488 0.00609 0.00132 -0.0459 0.4148 0.9889 3.500 0.6908 0.00622 0.00143 -0.0493 0.4028 0.9958 3.750 0.7339 0.00658 0.00158 -0.0533 0.3436 1.0000 4.000 0.7428 0.00868 0.00249 -0.0506 0.0768 1.0000 4.250 0.7645 0.00916 0.00281 -0.0496 0.0330 1.0000 4.500 0.7878 0.00943 0.00306 -0.0488 0.0210 1.0000 4.750 0.8106 0.00976 0.00336 -0.0480 0.0095 1.0000 5.000 0.8336 0.01006 0.00368 -0.0472 0.0069 1.0000 5.250 0.8563 0.01040 0.00406 -0.0463 0.0052 1.0000 5.500 0.8790 0.01073 0.00443 -0.0454 0.0043 1.0000 5.750 0.8999 0.01129 0.00506 -0.0442 0.0035 1.0000 6.000 0.9215 0.01173 0.00559 -0.0432 0.0032 1.0000 6.250 0.9420 0.01231 0.00626 -0.0419 0.0029 1.0000 6.500 0.9622 0.01288 0.00691 -0.0406 0.0026 1.0000 6.750 0.9834 0.01331 0.00737 -0.0397 0.0023 1.0000 7.000 1.0021 0.01401 0.00815 -0.0382 0.0020 1.0000 7.250 1.0191 0.01489 0.00914 -0.0365 0.0018 1.0000 7.500 1.0357 0.01581 0.01021 -0.0347 0.0018 1.0000 7.750 1.0517 0.01683 0.01135 -0.0328 0.0017 1.0000 8.000 1.0670 0.01797 0.01265 -0.0308 0.0015 1.0000 8.250 1.0806 0.01955 0.01441 -0.0285 0.0014 1.0000 8.500 1.0984 0.02048 0.01544 -0.0272 0.0011 1.0000 8.750 1.1122 0.02245 0.01764 -0.0251 0.0011 1.0000 9.250 1.1088 0.04067 0.03721 -0.0150 0.0012 1.0000 9.500 1.1064 0.04565 0.04250 -0.0115 0.0013 1.0000 9.750 1.1000 0.05025 0.04737 -0.0082 0.0013 1.0000 10.000 1.0920 0.05401 0.05133 -0.0053 0.0014 1.0000 10.250 1.0782 0.05722 0.05471 -0.0018 0.0014 1.0000 10.500 1.0584 0.06041 0.05804 0.0013 0.0014 1.0000 10.750 1.0410 0.06378 0.06154 0.0025 0.0014 1.0000 11.000 1.0233 0.06774 0.06562 0.0022 0.0015 1.0000 11.250 1.0061 0.07221 0.07020 0.0007 0.0015 1.0000 11.500 0.9869 0.07763 0.07573 -0.0019 0.0015 1.0000 11.750 0.9653 0.08427 0.08247 -0.0057 0.0015 1.0000 12.000 0.9517 0.09006 0.08834 -0.0095 0.0015 1.0000