XFOIL Version 6.96 Calculated polar for: USA 45 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 = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -15.250 -0.2583 0.14732 0.14575 -0.0210 1.0000 0.0152 -15.000 -0.2553 0.14579 0.14423 -0.0220 1.0000 0.0155 -14.750 -0.9293 0.04003 0.03728 -0.0799 1.0000 0.0151 -14.500 -0.9075 0.03995 0.03727 -0.0797 1.0000 0.0154 -14.250 -0.9055 0.03780 0.03506 -0.0804 1.0000 0.0155 -14.000 -0.9218 0.03443 0.03153 -0.0801 1.0000 0.0156 -13.750 -0.9145 0.03347 0.03053 -0.0789 1.0000 0.0158 -13.500 -0.9121 0.03232 0.02932 -0.0770 1.0000 0.0159 -13.250 -0.9121 0.03125 0.02818 -0.0743 1.0000 0.0161 -13.000 -0.9209 0.03002 0.02684 -0.0699 1.0000 0.0163 -12.750 -0.9283 0.02909 0.02582 -0.0649 1.0000 0.0164 -12.500 -0.9257 0.02827 0.02492 -0.0613 1.0000 0.0167 -12.250 -0.9262 0.02732 0.02386 -0.0572 1.0000 0.0169 -12.000 -0.9239 0.02659 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0.00282 -0.0415 0.3276 0.9729 2.500 0.6262 0.00814 0.00296 -0.0436 0.3233 0.9765 2.750 0.6597 0.00824 0.00306 -0.0448 0.3212 0.9811 3.000 0.6987 0.00835 0.00315 -0.0471 0.3182 0.9836 3.250 0.7392 0.00848 0.00324 -0.0499 0.3150 0.9856 3.500 0.7755 0.00862 0.00333 -0.0518 0.3117 0.9878 3.750 0.8090 0.00878 0.00345 -0.0531 0.3078 0.9901 4.000 0.8423 0.00886 0.00353 -0.0543 0.3059 0.9924 4.250 0.8764 0.00894 0.00361 -0.0557 0.3034 0.9942 4.500 0.9119 0.00902 0.00367 -0.0575 0.3004 0.9956 4.750 0.9467 0.00914 0.00375 -0.0592 0.2971 0.9973 5.000 0.9806 0.00931 0.00389 -0.0607 0.2929 0.9990 5.250 1.0139 0.00935 0.00394 -0.0620 0.2904 1.0000 5.500 1.0351 0.00941 0.00402 -0.0608 0.2875 1.0000 5.750 1.0557 0.00951 0.00411 -0.0594 0.2842 1.0000 6.000 1.0755 0.00966 0.00423 -0.0580 0.2805 1.0000 6.250 1.0959 0.00979 0.00436 -0.0566 0.2770 1.0000 6.500 1.1177 0.00986 0.00444 -0.0555 0.2736 1.0000 6.750 1.1386 0.00998 0.00456 -0.0542 0.2703 1.0000 7.000 1.1588 0.01014 0.00471 -0.0528 0.2666 1.0000 7.250 1.1790 0.01031 0.00488 -0.0514 0.2629 1.0000 7.500 1.2010 0.01040 0.00500 -0.0504 0.2596 1.0000 7.750 1.2220 0.01055 0.00514 -0.0492 0.2549 1.0000 8.000 1.2416 0.01077 0.00533 -0.0478 0.2496 1.0000 8.250 1.2634 0.01090 0.00548 -0.0467 0.2445 1.0000 8.500 1.2836 0.01109 0.00566 -0.0454 0.2388 1.0000 8.750 1.3036 0.01131 0.00587 -0.0441 0.2319 1.0000 9.000 1.3226 0.01158 0.00610 -0.0426 0.2233 1.0000 9.250 1.3398 0.01193 0.00639 -0.0409 0.2095 1.0000 9.500 1.3522 0.01252 0.00683 -0.0384 0.1859 1.0000 9.750 1.3576 0.01344 0.00754 -0.0349 0.1558 1.0000 10.000 1.3536 0.01458 0.00845 -0.0297 0.1231 1.0000 10.500 1.3198 0.01801 0.01141 -0.0165 0.0427 1.0000 10.750 1.3122 0.01965 0.01301 -0.0122 0.0204 1.0000 11.000 1.3205 0.02058 0.01398 -0.0102 0.0179 1.0000 11.250 1.3283 0.02162 0.01506 -0.0084 0.0167 1.0000 11.500 1.3359 0.02275 0.01626 -0.0068 0.0156 1.0000 11.750 1.3442 0.02393 0.01750 -0.0055 0.0150 1.0000 12.000 1.3511 0.02531 0.01894 -0.0044 0.0145 1.0000 12.250 1.3554 0.02699 0.02070 -0.0034 0.0139 1.0000 12.500 1.3588 0.02890 0.02268 -0.0026 0.0135 1.0000 12.750 1.3552 0.03163 0.02552 -0.0020 0.0129 1.0000 13.000 1.3539 0.03435 0.02834 -0.0019 0.0127 1.0000 13.250 1.3555 0.03689 0.03097 -0.0020 0.0125 1.0000 13.500 1.3529 0.04003 0.03420 -0.0024 0.0122 1.0000 13.750 1.3477 0.04357 0.03785 -0.0029 0.0122 1.0000 14.000 1.3393 0.04761 0.04199 -0.0037 0.0119 1.0000 14.250 1.3323 0.05153 0.04601 -0.0045 0.0116 1.0000 14.500 1.3186 0.05636 0.05096 -0.0055 0.0116 1.0000 14.750 1.3069 0.06099 0.05568 -0.0066 0.0114 1.0000 15.000 1.2933 0.06589 0.06068 -0.0077 0.0112 1.0000 15.250 1.2782 0.07104 0.06595 -0.0089 0.0113 1.0000 15.500 1.2646 0.07608 0.07108 -0.0102 0.0112 1.0000 15.750 1.2508 0.08120 0.07629 -0.0115 0.0111 1.0000 16.000 1.2385 0.08620 0.08138 -0.0129 0.0110 1.0000