XFOIL Version 6.96 Calculated polar for: NACA 63-415 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -18.250 -0.8075 0.12235 0.11935 -0.0302 1.0000 0.0130 -18.000 -0.8547 0.10829 0.10508 -0.0381 1.0000 0.0130 -17.750 -0.8829 0.09863 0.09526 -0.0436 1.0000 0.0130 -17.500 -0.9059 0.09028 0.08675 -0.0485 1.0000 0.0130 -17.250 -0.9249 0.08295 0.07928 -0.0527 1.0000 0.0131 -17.000 -0.9403 0.07656 0.07274 -0.0564 1.0000 0.0131 -16.750 -0.9529 0.07082 0.06685 -0.0597 1.0000 0.0132 -16.500 -0.9624 0.06581 0.06171 -0.0624 1.0000 0.0133 -16.250 -0.9698 0.06130 0.05706 -0.0648 1.0000 0.0134 -16.000 -0.9761 0.05714 0.05276 -0.0668 1.0000 0.0135 -15.750 -0.9821 0.05325 0.04878 -0.0687 1.0000 0.0137 -15.500 -0.9856 0.04985 0.04529 -0.0701 1.0000 0.0138 -15.250 -0.9878 0.04675 0.04213 -0.0713 1.0000 0.0140 -15.000 -0.9877 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0.8984 0.00993 0.00383 -0.0781 0.4038 0.7184 5.250 0.9216 0.01037 0.00411 -0.0775 0.3659 0.7255 5.500 0.9437 0.01087 0.00446 -0.0767 0.3244 0.7323 5.750 0.9646 0.01147 0.00488 -0.0758 0.2816 0.7398 6.000 0.9849 0.01212 0.00535 -0.0749 0.2404 0.7466 6.250 1.0047 0.01276 0.00584 -0.0738 0.2015 0.7542 6.500 1.0239 0.01343 0.00635 -0.0727 0.1670 0.7618 6.750 1.0429 0.01406 0.00687 -0.0715 0.1377 0.7696 7.000 1.0616 0.01469 0.00739 -0.0704 0.1120 0.7777 7.250 1.0799 0.01528 0.00792 -0.0690 0.0918 0.7863 7.500 1.0971 0.01590 0.00847 -0.0676 0.0735 0.7955 7.750 1.1133 0.01652 0.00904 -0.0660 0.0588 0.8052 8.000 1.1290 0.01709 0.00960 -0.0643 0.0487 0.8152 8.250 1.1427 0.01762 0.01014 -0.0622 0.0427 0.8267 8.500 1.1559 0.01814 0.01070 -0.0600 0.0389 0.8387 8.750 1.1689 0.01868 0.01130 -0.0579 0.0362 0.8516 9.000 1.1814 0.01923 0.01193 -0.0558 0.0340 0.8665 9.250 1.1917 0.01985 0.01263 -0.0534 0.0321 0.8851 9.500 1.2003 0.02046 0.01334 -0.0507 0.0308 0.9126 9.750 1.2120 0.02104 0.01404 -0.0489 0.0295 1.0000 10.000 1.2257 0.02189 0.01491 -0.0477 0.0280 1.0000 10.250 1.2378 0.02287 0.01591 -0.0464 0.0266 1.0000 10.500 1.2501 0.02387 0.01695 -0.0452 0.0257 1.0000 10.750 1.2632 0.02485 0.01798 -0.0442 0.0246 1.0000 11.000 1.2753 0.02593 0.01910 -0.0431 0.0234 1.0000 11.250 1.2860 0.02714 0.02035 -0.0421 0.0224 1.0000 11.500 1.2955 0.02849 0.02173 -0.0411 0.0214 1.0000 11.750 1.3070 0.02971 0.02303 -0.0402 0.0206 1.0000 12.000 1.3175 0.03105 0.02443 -0.0394 0.0196 1.0000 12.250 1.3268 0.03252 0.02594 -0.0387 0.0189 1.0000 12.500 1.3350 0.03414 0.02760 -0.0379 0.0181 1.0000 12.750 1.3424 0.03586 0.02938 -0.0372 0.0176 1.0000 13.000 1.3506 0.03755 0.03114 -0.0366 0.0170 1.0000 13.250 1.3580 0.03935 0.03302 -0.0361 0.0164 1.0000 13.500 1.3651 0.04124 0.03498 -0.0357 0.0159 1.0000 13.750 1.3710 0.04326 0.03706 -0.0353 0.0154 1.0000 14.000 1.3763 0.04543 0.03929 -0.0350 0.0151 1.0000 14.250 1.3803 0.04782 0.04173 -0.0349 0.0147 1.0000 14.500 1.3837 0.05032 0.04431 -0.0348 0.0144 1.0000 14.750 1.3877 0.05283 0.04691 -0.0349 0.0142 1.0000 15.000 1.3912 0.05548 0.04966 -0.0350 0.0139 1.0000 15.250 1.3941 0.05826 0.05253 -0.0353 0.0136 1.0000 15.500 1.3968 0.06114 0.05550 -0.0357 0.0134 1.0000 15.750 1.3990 0.06413 0.05858 -0.0362 0.0131 1.0000 16.000 1.4005 0.06732 0.06186 -0.0368 0.0129 1.0000 16.250 1.4016 0.07062 0.06525 -0.0376 0.0128 1.0000 16.500 1.4017 0.07413 0.06885 -0.0385 0.0126 1.0000 16.750 1.4019 0.07771 0.07252 -0.0396 0.0123 1.0000 17.000 1.4005 0.08160 0.07650 -0.0408 0.0122 1.0000 17.250 1.3980 0.08574 0.08073 -0.0422 0.0120 1.0000 17.500 1.3944 0.09013 0.08521 -0.0438 0.0119 1.0000 17.750 1.3902 0.09469 0.08987 -0.0456 0.0117 1.0000 18.000 1.3875 0.09911 0.09441 -0.0474 0.0116 1.0000 18.250 1.3845 0.10366 0.09908 -0.0493 0.0115 1.0000 18.500 1.3805 0.10843 0.10396 -0.0515 0.0114 1.0000 18.750 1.3762 0.11332 0.10896 -0.0537 0.0113 1.0000 19.000 1.3714 0.11836 0.11412 -0.0562 0.0112 1.0000 19.250 1.3665 0.12349 0.11936 -0.0588 0.0111 1.0000