XFOIL Version 6.96 Calculated polar for: FX 63-147 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 ------ -------- --------- --------- -------- -------- -------- -9.250 -0.4539 0.06447 0.06101 -0.0641 0.9969 0.0110 -9.000 -0.4612 0.05845 0.05479 -0.0698 0.9903 0.0110 -8.500 -0.4515 0.05326 0.04941 -0.0750 0.9733 0.0117 -8.250 -0.4515 0.04825 0.04411 -0.0764 0.9611 0.0117 -8.000 -0.4440 0.04533 0.04099 -0.0766 0.9497 0.0121 -7.750 -0.4366 0.04157 0.03692 -0.0764 0.9383 0.0123 -7.500 -0.4272 0.03693 0.03181 -0.0760 0.9281 0.0119 -7.250 -0.4140 0.03123 0.02517 -0.0750 0.9193 0.0111 -7.000 -0.3945 0.02898 0.02263 -0.0745 0.9091 0.0110 -6.750 -0.3655 0.02671 0.02005 -0.0755 0.9039 0.0109 -6.500 -0.3413 0.02490 0.01798 -0.0753 0.8938 0.0108 -6.250 -0.3079 0.02310 0.01590 -0.0767 0.8881 0.0108 -6.000 -0.2791 0.02166 0.01427 -0.0770 0.8778 0.0108 -5.750 -0.2442 0.02029 0.01273 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0.8244 0.01507 0.00742 -0.0805 0.4955 0.8536 4.250 0.8467 0.01519 0.00756 -0.0795 0.4892 0.8555 4.500 0.8699 0.01530 0.00772 -0.0788 0.4828 0.8577 4.750 0.8937 0.01542 0.00786 -0.0782 0.4748 0.8600 5.000 0.9180 0.01553 0.00801 -0.0778 0.4642 0.8625 5.250 0.9425 0.01567 0.00814 -0.0776 0.4505 0.8654 5.500 0.9693 0.01584 0.00831 -0.0779 0.4357 0.8683 5.750 0.9885 0.01598 0.00845 -0.0764 0.4242 0.8702 6.000 1.0085 0.01614 0.00864 -0.0751 0.4124 0.8723 6.250 1.0292 0.01633 0.00886 -0.0740 0.3988 0.8748 6.500 1.0507 0.01656 0.00909 -0.0732 0.3838 0.8774 6.750 1.0729 0.01683 0.00936 -0.0727 0.3686 0.8802 7.000 1.0956 0.01717 0.00968 -0.0723 0.3516 0.8827 7.250 1.1107 0.01747 0.00999 -0.0703 0.3347 0.8847 7.500 1.1284 0.01783 0.01035 -0.0689 0.3169 0.8867 7.750 1.1469 0.01824 0.01077 -0.0677 0.3001 0.8889 8.000 1.1648 0.01873 0.01126 -0.0665 0.2822 0.8916 8.250 1.1825 0.01935 0.01185 -0.0655 0.2599 0.8941 8.500 1.1988 0.02016 0.01258 -0.0645 0.2333 0.8964 8.750 1.2102 0.02097 0.01333 -0.0625 0.2066 0.8984 9.000 1.2166 0.02199 0.01424 -0.0599 0.1788 0.9005 9.250 1.2243 0.02304 0.01524 -0.0576 0.1536 0.9030 9.500 1.2301 0.02433 0.01645 -0.0554 0.1302 0.9058 9.750 1.2392 0.02558 0.01768 -0.0538 0.1088 0.9082 10.000 1.2470 0.02706 0.01913 -0.0525 0.0929 0.9103 10.250 1.2528 0.02851 0.02060 -0.0509 0.0787 0.9122 10.500 1.2538 0.03021 0.02228 -0.0487 0.0619 0.9143 10.750 1.2561 0.03198 0.02408 -0.0471 0.0537 0.9167 11.000 1.2610 0.03374 0.02591 -0.0460 0.0435 0.9191 11.250 1.2637 0.03588 0.02808 -0.0451 0.0349 0.9212 11.500 1.2665 0.03820 0.03044 -0.0446 0.0313 0.9231 11.750 1.2702 0.04059 0.03291 -0.0444 0.0286 0.9247 12.000 1.2745 0.04255 0.03501 -0.0437 0.0267 0.9267 12.250 1.2773 0.04468 0.03728 -0.0430 0.0245 0.9289 12.500 1.2792 0.04710 0.03982 -0.0426 0.0227 0.9310 12.750 1.2805 0.04972 0.04257 -0.0425 0.0214 0.9329 13.000 1.2837 0.05230 0.04529 -0.0426 0.0203 0.9346 13.250 1.2863 0.05505 0.04817 -0.0429 0.0191 0.9362 13.500 1.2877 0.05798 0.05121 -0.0433 0.0179 0.9377 13.750 1.2874 0.06107 0.05439 -0.0436 0.0169 0.9392 14.000 1.2855 0.06410 0.05754 -0.0436 0.0163 0.9411 14.250 1.2836 0.06724 0.06081 -0.0438 0.0159 0.9433 14.500 1.2816 0.07051 0.06420 -0.0441 0.0154 0.9453 14.750 1.2798 0.07399 0.06780 -0.0448 0.0150 0.9471 15.000 1.2773 0.07760 0.07154 -0.0456 0.0146 0.9488 15.250 1.2746 0.08142 0.07549 -0.0465 0.0143 0.9503 15.750 1.2652 0.08929 0.08361 -0.0483 0.0136 0.9536 16.000 1.2587 0.09346 0.08789 -0.0492 0.0132 0.9554