XFOIL Version 6.96 Calculated polar for: GOE 645 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 ------ -------- --------- --------- -------- -------- -------- -17.000 -0.8284 0.08561 0.08263 -0.0611 1.0000 0.0173 -16.750 -0.8484 0.07930 0.07626 -0.0647 1.0000 0.0175 -16.500 -0.8691 0.07311 0.06999 -0.0682 1.0000 0.0176 -16.250 -0.8966 0.06634 0.06314 -0.0722 1.0000 0.0178 -16.000 -0.9339 0.05882 0.05551 -0.0767 1.0000 0.0177 -15.750 -0.9794 0.05071 0.04729 -0.0815 1.0000 0.0176 -15.500 -1.1123 0.03007 0.02622 -0.0979 1.0000 0.0168 -15.250 -1.1356 0.02752 0.02354 -0.0947 1.0000 0.0169 -15.000 -1.1219 0.02597 0.02190 -0.0950 0.9991 0.0173 -14.750 -1.0980 0.02448 0.02032 -0.0968 0.9967 0.0176 -14.500 -1.0727 0.02328 0.01907 -0.0982 0.9939 0.0181 -14.250 -1.0491 0.02233 0.01806 -0.0988 0.9907 0.0187 -14.000 -1.0235 0.02144 0.01712 -0.0995 0.9871 0.0192 -13.750 -0.9953 0.02064 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0.6465 0.00959 0.00235 -0.0857 0.3114 0.2769 1.750 0.6722 0.00973 0.00243 -0.0852 0.2995 0.2815 2.000 0.6976 0.00986 0.00253 -0.0846 0.2877 0.2877 2.250 0.7228 0.01001 0.00264 -0.0840 0.2768 0.2949 2.500 0.7476 0.01019 0.00277 -0.0834 0.2647 0.3005 2.750 0.7727 0.01032 0.00288 -0.0828 0.2552 0.3081 3.000 0.7972 0.01050 0.00302 -0.0820 0.2442 0.3154 3.250 0.8212 0.01069 0.00318 -0.0813 0.2336 0.3243 3.500 0.8464 0.01081 0.00330 -0.0807 0.2279 0.3335 3.750 0.8709 0.01095 0.00344 -0.0800 0.2215 0.3435 4.250 0.9200 0.01121 0.00373 -0.0786 0.2121 0.3693 4.500 0.9441 0.01133 0.00388 -0.0779 0.2075 0.3885 4.750 0.9669 0.01146 0.00406 -0.0770 0.2017 0.4243 5.000 0.9902 0.01152 0.00424 -0.0761 0.1977 0.4696 5.250 1.0122 0.01157 0.00443 -0.0751 0.1922 0.5408 5.500 1.0329 0.01166 0.00465 -0.0737 0.1862 0.6088 7.000 1.1779 0.01342 0.00680 -0.0714 0.1105 1.0000 7.250 1.1961 0.01366 0.00704 -0.0696 0.1096 1.0000 7.500 1.2142 0.01390 0.00730 -0.0679 0.1089 1.0000 7.750 1.2325 0.01417 0.00757 -0.0662 0.1083 1.0000 8.000 1.2504 0.01446 0.00787 -0.0645 0.1075 1.0000 8.250 1.2687 0.01475 0.00818 -0.0630 0.1071 1.0000 8.500 1.2867 0.01507 0.00851 -0.0614 0.1066 1.0000 8.750 1.3045 0.01541 0.00887 -0.0599 0.1061 1.0000 9.000 1.3219 0.01578 0.00925 -0.0584 0.1055 1.0000 9.250 1.3393 0.01616 0.00966 -0.0569 0.1051 1.0000 9.500 1.3566 0.01657 0.01009 -0.0555 0.1048 1.0000 9.750 1.3731 0.01703 0.01057 -0.0541 0.1043 1.0000 10.000 1.3888 0.01755 0.01111 -0.0526 0.1033 1.0000 10.250 1.4044 0.01810 0.01169 -0.0512 0.1024 1.0000 10.500 1.4184 0.01877 0.01237 -0.0497 0.1007 1.0000 10.750 1.4331 0.01943 0.01308 -0.0484 0.0998 1.0000 11.000 1.4475 0.02016 0.01384 -0.0472 0.0991 1.0000 11.250 1.4613 0.02095 0.01467 -0.0460 0.0984 1.0000 11.500 1.4768 0.02168 0.01544 -0.0450 0.0982 1.0000 11.750 1.4924 0.02242 0.01623 -0.0442 0.0979 1.0000 12.000 1.5073 0.02324 0.01709 -0.0434 0.0977 1.0000 12.250 1.5221 0.02409 0.01799 -0.0426 0.0972 1.0000 12.500 1.5360 0.02505 0.01900 -0.0419 0.0969 1.0000 12.750 1.5499 0.02602 0.02002 -0.0412 0.0960 1.0000 13.000 1.5629 0.02709 0.02114 -0.0405 0.0950 1.0000 13.250 1.5746 0.02828 0.02236 -0.0399 0.0933 1.0000 13.500 1.5848 0.02964 0.02377 -0.0392 0.0923 1.0000 13.750 1.5934 0.03115 0.02531 -0.0385 0.0905 1.0000 14.000 1.6012 0.03275 0.02696 -0.0379 0.0896 1.0000 14.250 1.6072 0.03455 0.02881 -0.0373 0.0884 1.0000 14.500 1.6141 0.03632 0.03063 -0.0368 0.0869 1.0000 14.750 1.6242 0.03779 0.03215 -0.0364 0.0855 1.0000 15.000 1.6315 0.03954 0.03397 -0.0360 0.0831 1.0000 15.250 1.6322 0.04199 0.03640 -0.0356 0.0756 1.0000 15.500 1.6109 0.04680 0.04110 -0.0352 0.0571 1.0000 15.750 1.5857 0.05232 0.04663 -0.0351 0.0476 1.0000 16.000 1.5705 0.05693 0.05130 -0.0353 0.0424 1.0000 16.250 1.5571 0.06141 0.05584 -0.0356 0.0383 1.0000 16.500 1.5419 0.06621 0.06071 -0.0361 0.0339 1.0000 16.750 1.5190 0.07206 0.06661 -0.0368 0.0274 1.0000 17.000 1.4976 0.07785 0.07244 -0.0377 0.0216 1.0000 17.250 1.4789 0.08340 0.07804 -0.0387 0.0177 1.0000 17.500 1.4664 0.08815 0.08287 -0.0396 0.0161 1.0000