XFOIL Version 6.96 Calculated polar for: GOE 735 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -16.750 -0.7578 0.08302 0.07843 -0.0518 1.0000 0.0527 -16.500 -0.8029 0.07235 0.06758 -0.0590 1.0000 0.0531 -16.250 -0.8338 0.06418 0.05927 -0.0647 1.0000 0.0536 -16.000 -0.8555 0.05771 0.05267 -0.0692 1.0000 0.0541 -15.500 -0.8750 0.04788 0.04249 -0.0774 0.9115 0.0554 -15.250 -0.8742 0.04421 0.03852 -0.0809 0.8735 0.0564 -15.000 -0.8799 0.04126 0.03531 -0.0816 0.8509 0.0574 -14.750 -0.8845 0.03881 0.03260 -0.0814 0.8351 0.0583 -14.500 -0.8831 0.03719 0.03090 -0.0806 0.8220 0.0590 -14.250 -0.8812 0.03579 0.02940 -0.0794 0.8103 0.0599 -14.000 -0.8790 0.03454 0.02803 -0.0778 0.8002 0.0609 -13.750 -0.8760 0.03334 0.02669 -0.0760 0.7907 0.0621 -13.500 -0.8723 0.03215 0.02531 -0.0740 0.7823 0.0635 -13.250 -0.8666 0.03109 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0.5021 0.01514 0.00746 -0.0217 0.4000 0.9020 2.000 0.5385 0.01551 0.00778 -0.0230 0.3964 0.9191 2.250 0.5773 0.01592 0.00810 -0.0248 0.3929 0.9345 2.500 0.6186 0.01628 0.00836 -0.0272 0.3896 0.9440 2.750 0.6532 0.01652 0.00861 -0.0286 0.3860 0.9512 3.000 0.6931 0.01676 0.00883 -0.0311 0.3825 0.9573 3.250 0.7264 0.01702 0.00906 -0.0323 0.3793 0.9652 3.500 0.7665 0.01722 0.00919 -0.0349 0.3763 0.9695 3.750 0.7990 0.01743 0.00931 -0.0361 0.3736 0.9740 4.000 0.8269 0.01766 0.00952 -0.0364 0.3708 0.9787 4.250 0.8624 0.01782 0.00971 -0.0383 0.3674 0.9816 4.500 0.8980 0.01800 0.00990 -0.0402 0.3640 0.9856 4.750 0.9318 0.01822 0.01009 -0.0418 0.3612 0.9900 5.000 0.9654 0.01840 0.01024 -0.0434 0.3585 0.9934 5.250 0.9978 0.01859 0.01037 -0.0447 0.3562 0.9962 5.500 1.0294 0.01882 0.01057 -0.0460 0.3537 0.9989 5.750 1.0529 0.01906 0.01088 -0.0457 0.3509 1.0000 6.000 1.0693 0.01933 0.01118 -0.0441 0.3481 1.0000 6.250 1.0855 0.01960 0.01146 -0.0424 0.3456 1.0000 6.500 1.1013 0.01987 0.01175 -0.0407 0.3433 1.0000 6.750 1.1170 0.02016 0.01201 -0.0389 0.3412 1.0000 7.000 1.1329 0.02045 0.01226 -0.0372 0.3391 1.0000 7.250 1.1480 0.02077 0.01256 -0.0353 0.3369 1.0000 7.500 1.1590 0.02115 0.01302 -0.0329 0.3345 1.0000 7.750 1.1693 0.02154 0.01346 -0.0304 0.3321 1.0000 8.000 1.1788 0.02193 0.01388 -0.0277 0.3297 1.0000 8.250 1.1872 0.02231 0.01429 -0.0249 0.3273 1.0000 8.500 1.1926 0.02267 0.01465 -0.0216 0.3251 1.0000 8.750 1.1975 0.02301 0.01497 -0.0182 0.3230 1.0000 9.000 1.2074 0.02338 0.01529 -0.0156 0.3211 1.0000 9.250 1.2080 0.02396 0.01594 -0.0121 0.3189 1.0000 9.500 1.2084 0.02467 0.01674 -0.0089 0.3165 1.0000 9.750 1.2114 0.02542 0.01755 -0.0062 0.3139 1.0000 10.000 1.2168 0.02618 0.01835 -0.0039 0.3114 1.0000 10.250 1.2241 0.02692 0.01911 -0.0020 0.3090 1.0000 10.500 1.2350 0.02756 0.01974 -0.0005 0.3067 1.0000 10.750 1.2500 0.02809 0.02023 0.0007 0.3048 1.0000 11.000 1.2554 0.02912 0.02132 0.0024 0.3026 1.0000 11.250 1.2556 0.03049 0.02281 0.0042 0.3000 1.0000 11.500 1.2579 0.03183 0.02424 0.0057 0.2973 1.0000 11.750 1.2619 0.03313 0.02559 0.0070 0.2946 1.0000 12.000 1.2692 0.03425 0.02674 0.0081 0.2920 1.0000 12.250 1.2821 0.03503 0.02751 0.0089 0.2899 1.0000 12.500 1.2991 0.03558 0.02802 0.0096 0.2878 1.0000 12.750 1.2877 0.03821 0.03083 0.0109 0.2849 1.0000 13.000 1.2808 0.04067 0.03342 0.0119 0.2818 1.0000 13.250 1.2782 0.04286 0.03569 0.0127 0.2789 1.0000 13.500 1.2819 0.04458 0.03745 0.0132 0.2762 1.0000 13.750 1.2961 0.04540 0.03825 0.0136 0.2741 1.0000 14.000 1.3077 0.04646 0.03932 0.0141 0.2716 1.0000 14.250 1.2682 0.05265 0.04576 0.0142 0.2668 1.0000 14.500 1.2532 0.05667 0.04989 0.0141 0.2628 1.0000 14.750 1.2591 0.05850 0.05174 0.0142 0.2601 1.0000 15.000 1.2786 0.05882 0.05204 0.0145 0.2583 1.0000 15.250 1.1907 0.07176 0.06529 0.0126 0.2497 1.0000 15.500 1.1866 0.07501 0.06858 0.0122 0.2461 1.0000 15.750 1.2065 0.07526 0.06883 0.0124 0.2443 1.0000 16.000 1.2343 0.07454 0.06808 0.0128 0.2430 1.0000 16.500 1.1056 0.09725 0.09111 0.0080 0.2256 1.0000 16.750 1.1499 0.09408 0.08791 0.0089 0.2267 1.0000