XFOIL Version 6.96 Calculated polar for: CLARK W 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 = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -14.750 -0.9347 0.05878 0.05633 -0.0701 1.0000 0.0180 -14.500 -1.0485 0.03472 0.03179 -0.0919 1.0000 0.0176 -14.250 -1.0624 0.03272 0.02968 -0.0883 1.0000 0.0177 -14.000 -1.0569 0.03094 0.02777 -0.0874 0.9990 0.0179 -13.750 -1.0371 0.02921 0.02589 -0.0887 0.9967 0.0182 -13.500 -1.0145 0.02779 0.02433 -0.0899 0.9949 0.0184 -13.250 -0.9909 0.02658 0.02299 -0.0908 0.9932 0.0186 -13.000 -0.9679 0.02553 0.02183 -0.0913 0.9905 0.0188 -12.750 -0.9431 0.02454 0.02071 -0.0921 0.9883 0.0190 -12.500 -0.9183 0.02340 0.01944 -0.0929 0.9864 0.0191 -12.250 -0.8941 0.02211 0.01805 -0.0937 0.9844 0.0195 -12.000 -0.8695 0.02136 0.01724 -0.0939 0.9814 0.0197 -11.750 -0.8446 0.02071 0.01653 -0.0940 0.9778 0.0199 -11.500 -0.8184 0.02004 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1.1707 0.01252 0.00612 -0.0672 0.0935 1.0000 8.000 1.1911 0.01296 0.00649 -0.0660 0.0800 1.0000 8.250 1.2108 0.01344 0.00689 -0.0647 0.0656 1.0000 8.500 1.2303 0.01392 0.00730 -0.0634 0.0549 1.0000 8.750 1.2503 0.01436 0.00771 -0.0622 0.0480 1.0000 9.000 1.2695 0.01480 0.00812 -0.0608 0.0428 1.0000 9.250 1.2885 0.01518 0.00850 -0.0594 0.0395 1.0000 9.500 1.3065 0.01562 0.00894 -0.0578 0.0364 1.0000 9.750 1.3254 0.01601 0.00934 -0.0564 0.0346 1.0000 10.000 1.3434 0.01646 0.00980 -0.0550 0.0325 1.0000 10.250 1.3606 0.01696 0.01030 -0.0534 0.0304 1.0000 10.500 1.3787 0.01740 0.01077 -0.0520 0.0293 1.0000 10.750 1.3961 0.01788 0.01128 -0.0506 0.0279 1.0000 11.000 1.4123 0.01844 0.01186 -0.0491 0.0263 1.0000 11.250 1.4281 0.01904 0.01247 -0.0475 0.0249 1.0000 11.500 1.4446 0.01959 0.01306 -0.0461 0.0238 1.0000 11.750 1.4600 0.02021 0.01371 -0.0446 0.0225 1.0000 12.000 1.4742 0.02093 0.01445 -0.0430 0.0212 1.0000 12.250 1.4885 0.02165 0.01521 -0.0415 0.0201 1.0000 12.500 1.5024 0.02241 0.01600 -0.0400 0.0189 1.0000 12.750 1.5149 0.02328 0.01690 -0.0385 0.0176 1.0000 13.000 1.5268 0.02421 0.01786 -0.0370 0.0166 1.0000 13.250 1.5388 0.02516 0.01886 -0.0355 0.0157 1.0000 13.500 1.5498 0.02622 0.01996 -0.0341 0.0149 1.0000 13.750 1.5594 0.02740 0.02118 -0.0326 0.0142 1.0000 14.000 1.5682 0.02867 0.02251 -0.0312 0.0135 1.0000 14.250 1.5776 0.02995 0.02385 -0.0299 0.0131 1.0000 14.500 1.5858 0.03135 0.02531 -0.0287 0.0126 1.0000 14.750 1.5928 0.03289 0.02691 -0.0275 0.0121 1.0000 15.000 1.5985 0.03459 0.02867 -0.0264 0.0116 1.0000 15.250 1.6028 0.03648 0.03062 -0.0253 0.0111 1.0000 15.500 1.6061 0.03851 0.03272 -0.0243 0.0107 1.0000 15.750 1.6105 0.04051 0.03480 -0.0235 0.0105 1.0000 16.000 1.6138 0.04270 0.03707 -0.0229 0.0102 1.0000 16.250 1.6159 0.04509 0.03954 -0.0224 0.0100 1.0000 16.500 1.6166 0.04770 0.04224 -0.0220 0.0097 1.0000 16.750 1.6162 0.05054 0.04516 -0.0218 0.0095 1.0000 17.000 1.6147 0.05361 0.04831 -0.0217 0.0093 1.0000 17.250 1.6116 0.05695 0.05174 -0.0219 0.0091 1.0000 17.500 1.6069 0.06058 0.05546 -0.0222 0.0089 1.0000 17.750 1.6009 0.06448 0.05946 -0.0227 0.0087 1.0000 18.000 1.5930 0.06877 0.06385 -0.0235 0.0085 1.0000 18.250 1.5842 0.07326 0.06844 -0.0245 0.0084 1.0000 18.500 1.5761 0.07773 0.07302 -0.0256 0.0083 1.0000 18.750 1.5668 0.08245 0.07785 -0.0268 0.0082 1.0000 19.000 1.5569 0.08734 0.08285 -0.0282 0.0081 1.0000 19.250 1.5452 0.09255 0.08817 -0.0299 0.0081 1.0000