XFOIL Version 6.96 Calculated polar for: NACA 65(4)-421 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 ------ -------- --------- --------- -------- -------- -------- -19.750 -0.8928 0.08986 0.08551 -0.0788 1.0000 0.0209 -19.500 -0.8965 0.08599 0.08155 -0.0805 1.0000 0.0211 -19.250 -0.8989 0.08236 0.07783 -0.0822 1.0000 0.0212 -19.000 -0.9010 0.07889 0.07426 -0.0836 1.0000 0.0213 -18.750 -0.9023 0.07562 0.07090 -0.0848 1.0000 0.0215 -18.500 -0.9022 0.07260 0.06779 -0.0858 1.0000 0.0216 -18.250 -0.9016 0.06967 0.06476 -0.0867 1.0000 0.0218 -18.000 -0.8999 0.06697 0.06197 -0.0875 1.0000 0.0220 -17.750 -0.8974 0.06442 0.05933 -0.0881 1.0000 0.0222 -17.500 -0.8941 0.06205 0.05688 -0.0886 1.0000 0.0223 -17.250 -0.8898 0.05981 0.05455 -0.0890 1.0000 0.0225 -17.000 -0.8858 0.05754 0.05218 -0.0895 1.0000 0.0228 -16.750 -0.8664 0.05509 0.04963 -0.0924 0.9686 0.0228 -16.500 -0.8413 0.05250 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0.6323 3.250 0.6940 0.01009 0.00392 -0.0774 0.5835 0.6346 3.500 0.7223 0.01015 0.00401 -0.0777 0.5816 0.6372 3.750 0.7508 0.01024 0.00411 -0.0781 0.5796 0.6399 4.000 0.7787 0.01032 0.00423 -0.0784 0.5774 0.6426 4.250 0.8057 0.01038 0.00435 -0.0785 0.5745 0.6451 4.500 0.8325 0.01043 0.00445 -0.0786 0.5708 0.6473 4.750 0.8589 0.01048 0.00454 -0.0786 0.5670 0.6498 5.000 0.8853 0.01054 0.00462 -0.0785 0.5627 0.6523 5.250 0.9091 0.01058 0.00474 -0.0780 0.5555 0.6549 5.500 0.9318 0.01064 0.00479 -0.0772 0.5463 0.6578 5.750 0.9545 0.01071 0.00492 -0.0765 0.5376 0.6609 6.000 0.9759 0.01083 0.00502 -0.0755 0.5287 0.6641 6.250 0.9961 0.01094 0.00517 -0.0744 0.5171 0.6671 6.500 1.0098 0.01111 0.00533 -0.0719 0.5000 0.6702 6.750 1.0160 0.01137 0.00555 -0.0681 0.4794 0.6735 7.000 1.0223 0.01180 0.00593 -0.0646 0.4552 0.6769 7.250 1.0182 0.01260 0.00657 -0.0595 0.4206 0.6806 7.500 1.0083 0.01375 0.00753 -0.0540 0.3827 0.6841 7.750 0.9964 0.01517 0.00878 -0.0486 0.3449 0.6876 8.000 0.9871 0.01670 0.01017 -0.0440 0.3098 0.6914 8.250 0.9791 0.01834 0.01167 -0.0400 0.2759 0.6956 8.500 0.9714 0.02011 0.01330 -0.0363 0.2425 0.7001 8.750 0.9663 0.02189 0.01493 -0.0332 0.2121 0.7042 9.000 0.9622 0.02368 0.01660 -0.0303 0.1830 0.7083 9.250 0.9637 0.02523 0.01809 -0.0282 0.1610 0.7127 9.500 0.9659 0.02680 0.01958 -0.0263 0.1402 0.7172 9.750 0.9682 0.02846 0.02115 -0.0245 0.1196 0.7216 10.000 0.9708 0.03014 0.02275 -0.0228 0.0998 0.7258 10.250 0.9739 0.03182 0.02436 -0.0213 0.0813 0.7303 10.500 0.9784 0.03348 0.02597 -0.0200 0.0663 0.7353 10.750 0.9833 0.03516 0.02759 -0.0188 0.0515 0.7406 11.000 0.9893 0.03680 0.02920 -0.0177 0.0410 0.7455 11.250 0.9971 0.03832 0.03072 -0.0168 0.0349 0.7511 11.500 1.0068 0.03975 0.03219 -0.0161 0.0316 0.7570 11.750 1.0175 0.04114 0.03360 -0.0155 0.0294 0.7626 12.000 1.0279 0.04254 0.03506 -0.0149 0.0276 0.7681 12.250 1.0376 0.04403 0.03660 -0.0144 0.0262 0.7742 12.500 1.0489 0.04543 0.03805 -0.0140 0.0251 0.7805 12.750 1.0594 0.04691 0.03961 -0.0135 0.0242 0.7866 13.000 1.0691 0.04846 0.04123 -0.0131 0.0233 0.7939 13.250 1.0786 0.05009 0.04290 -0.0127 0.0226 0.8012 13.500 1.0873 0.05178 0.04467 -0.0123 0.0219 0.8087 13.750 1.0975 0.05337 0.04634 -0.0120 0.0214 0.8170 14.000 1.1072 0.05501 0.04806 -0.0117 0.0208 0.8248 14.250 1.1159 0.05679 0.04991 -0.0115 0.0201 0.8335 14.500 1.1240 0.05864 0.05183 -0.0112 0.0194 0.8420 14.750 1.1311 0.06059 0.05385 -0.0110 0.0191 0.8517 15.000 1.1369 0.06268 0.05602 -0.0107 0.0188 0.8621 15.250 1.1448 0.06451 0.05797 -0.0105 0.0184 0.8749 15.500 1.1509 0.06648 0.06005 -0.0101 0.0182 0.8915 15.750 1.1568 0.06850 0.06219 -0.0098 0.0179 0.9169 16.000 1.1689 0.07084 0.06466 -0.0113 0.0175 0.9590 16.250 1.1734 0.07312 0.06700 -0.0113 0.0172 1.0000 16.500 1.1803 0.07547 0.06940 -0.0117 0.0170 1.0000 16.750 1.1861 0.07798 0.07197 -0.0122 0.0167 1.0000 17.000 1.1914 0.08059 0.07463 -0.0127 0.0165 1.0000 17.250 1.1957 0.08336 0.07745 -0.0133 0.0163 1.0000 17.500 1.1988 0.08630 0.08045 -0.0140 0.0161 1.0000 17.750 1.2008 0.08941 0.08361 -0.0147 0.0158 1.0000