XFOIL Version 6.94 Calculated polar for: NYU/GRUMMAN K-1 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.3026 0.01125 0.00616 -0.0902 0.8906 0.6703 0.500 0.3542 0.01073 0.00572 -0.0877 0.8734 0.6810 1.000 0.4084 0.01044 0.00548 -0.0862 0.8569 0.6927 1.500 0.4634 0.01013 0.00525 -0.0850 0.8393 0.7032 2.000 0.5168 0.00988 0.00509 -0.0834 0.8128 0.7132 2.500 0.5721 0.00969 0.00489 -0.0822 0.7715 0.7241 3.000 0.6211 0.00977 0.00460 -0.0795 0.6503 0.7333 3.500 0.6509 0.01254 0.00551 -0.0752 0.2828 0.7434 4.000 0.7002 0.01371 0.00620 -0.0745 0.1988 0.7531 4.500 0.7503 0.01438 0.00679 -0.0735 0.1699 0.7629 5.000 0.8022 0.01506 0.00740 -0.0729 0.1477 0.7730 5.500 0.8549 0.01556 0.00794 -0.0724 0.1265 0.7831 6.000 0.9060 0.01613 0.00848 -0.0716 0.0901 0.7935 6.500 0.9527 0.01764 0.00965 -0.0703 0.0505 0.8036 7.000 0.9983 0.01883 0.01089 -0.0686 0.0408 0.8151 8.000 1.0865 0.02170 0.01390 -0.0651 0.0325 0.8389 8.500 1.1283 0.02343 0.01577 -0.0630 0.0297 0.8541 9.000 1.1674 0.02534 0.01783 -0.0605 0.0275 0.8717 9.500 1.2046 0.02722 0.01994 -0.0576 0.0256 0.8994 10.500 1.2774 0.03244 0.02568 -0.0528 0.0227 1.0000 11.000 1.3077 0.03619 0.02976 -0.0502 0.0217 1.0000 11.500 1.3179 0.04054 0.03470 -0.0451 0.0209 1.0000 12.000 1.3177 0.04520 0.03981 -0.0399 0.0203 1.0000 12.500 1.3115 0.05034 0.04531 -0.0354 0.0199 1.0000 13.000 1.2971 0.05658 0.05190 -0.0320 0.0197 1.0000 13.500 1.2542 0.06646 0.06237 -0.0310 0.0197 1.0000 14.000 1.1464 0.09069 0.08757 -0.0440 0.0205 1.0000 14.500 1.0170 0.13560 0.13304 -0.0755 0.0223 1.0000