XFOIL Version 6.96 Calculated polar for: NPL 9510 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 ------ -------- --------- --------- -------- -------- -------- -19.750 -1.1789 0.12410 0.12100 -0.0096 1.0000 0.0126 -19.500 -1.1973 0.11685 0.11365 -0.0133 1.0000 0.0127 -19.250 -1.2156 0.10976 0.10646 -0.0169 1.0000 0.0127 -19.000 -1.2325 0.10302 0.09961 -0.0203 1.0000 0.0128 -18.750 -1.2491 0.09648 0.09296 -0.0235 1.0000 0.0128 -18.500 -1.2633 0.09048 0.08686 -0.0264 1.0000 0.0129 -18.250 -1.2759 0.08481 0.08109 -0.0291 1.0000 0.0129 -18.000 -1.2870 0.07951 0.07568 -0.0316 1.0000 0.0130 -17.750 -1.2962 0.07459 0.07067 -0.0339 1.0000 0.0131 -17.500 -1.3034 0.07008 0.06607 -0.0359 1.0000 0.0132 -17.250 -1.3092 0.06588 0.06177 -0.0378 1.0000 0.0133 -17.000 -1.3131 0.06198 0.05779 -0.0395 1.0000 0.0134 -16.750 -1.3156 0.05838 0.05410 -0.0410 1.0000 0.0135 -16.500 -1.3171 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0.0282 0.5393 2.750 0.5405 0.01071 0.00336 -0.0853 0.0273 0.5426 3.000 0.5687 0.01087 0.00354 -0.0852 0.0265 0.5460 3.250 0.5967 0.01105 0.00373 -0.0851 0.0257 0.5492 3.500 0.6245 0.01124 0.00392 -0.0849 0.0250 0.5524 3.750 0.6522 0.01146 0.00414 -0.0848 0.0243 0.5555 4.000 0.6795 0.01174 0.00442 -0.0845 0.0235 0.5587 4.250 0.7073 0.01192 0.00464 -0.0844 0.0229 0.5622 4.500 0.7350 0.01213 0.00487 -0.0842 0.0223 0.5656 4.750 0.7624 0.01235 0.00512 -0.0840 0.0217 0.5690 5.000 0.7898 0.01259 0.00536 -0.0838 0.0212 0.5722 5.250 0.8170 0.01284 0.00562 -0.0835 0.0206 0.5754 5.500 0.8439 0.01314 0.00594 -0.0832 0.0200 0.5792 5.750 0.8707 0.01345 0.00629 -0.0829 0.0196 0.5830 6.000 0.8976 0.01373 0.00661 -0.0826 0.0193 0.5865 6.250 0.9243 0.01403 0.00694 -0.0822 0.0189 0.5899 6.500 0.9510 0.01432 0.00727 -0.0819 0.0184 0.5933 6.750 0.9777 0.01460 0.00759 -0.0816 0.0179 0.5975 7.000 1.0043 0.01488 0.00789 -0.0812 0.0175 0.6016 7.250 1.0306 0.01519 0.00822 -0.0808 0.0171 0.6055 7.500 1.0562 0.01562 0.00868 -0.0803 0.0167 0.6089 7.750 1.0819 0.01602 0.00914 -0.0798 0.0164 0.6131 8.000 1.1077 0.01641 0.00960 -0.0793 0.0162 0.6177 8.250 1.1331 0.01683 0.01008 -0.0788 0.0159 0.6222 8.500 1.1584 0.01726 0.01057 -0.0783 0.0156 0.6260 8.750 1.1835 0.01769 0.01107 -0.0777 0.0154 0.6303 9.000 1.2084 0.01813 0.01157 -0.0771 0.0151 0.6353 9.250 1.2332 0.01857 0.01208 -0.0765 0.0148 0.6404 9.500 1.2577 0.01904 0.01260 -0.0758 0.0146 0.6450 9.750 1.2819 0.01951 0.01314 -0.0751 0.0144 0.6501 10.000 1.3058 0.02002 0.01371 -0.0744 0.0142 0.6555 10.250 1.3290 0.02062 0.01437 -0.0736 0.0140 0.6610 10.500 1.3510 0.02135 0.01520 -0.0726 0.0137 0.6672 10.750 1.3729 0.02208 0.01603 -0.0716 0.0136 0.6738 11.000 1.3948 0.02276 0.01681 -0.0706 0.0135 0.6798 11.250 1.4160 0.02349 0.01766 -0.0695 0.0134 0.6864 11.500 1.4366 0.02427 0.01856 -0.0683 0.0133 0.6935 11.750 1.4566 0.02507 0.01948 -0.0671 0.0131 0.7005 12.000 1.4761 0.02589 0.02042 -0.0658 0.0130 0.7082 12.250 1.4949 0.02674 0.02138 -0.0644 0.0128 0.7149 12.500 1.5131 0.02759 0.02236 -0.0630 0.0127 0.7221 12.750 1.5302 0.02851 0.02339 -0.0614 0.0125 0.7290 13.000 1.5462 0.02946 0.02447 -0.0597 0.0124 0.7362 13.250 1.5609 0.03049 0.02563 -0.0578 0.0123 0.7444 13.500 1.5742 0.03156 0.02682 -0.0558 0.0122 0.7530 14.000 1.5943 0.03386 0.02941 -0.0509 0.0120 0.7728 14.250 1.5975 0.03506 0.03074 -0.0474 0.0119 0.7843 14.500 1.6003 0.03644 0.03227 -0.0442 0.0118 0.7986 14.750 1.6016 0.03798 0.03398 -0.0410 0.0118 0.8169 15.000 1.6010 0.03972 0.03590 -0.0380 0.0117 0.8432 15.250 1.5951 0.04139 0.03790 -0.0344 0.0117 1.0000 15.500 1.5904 0.04374 0.04038 -0.0319 0.0116 1.0000 15.750 1.5825 0.04655 0.04334 -0.0297 0.0115 1.0000 16.000 1.5704 0.05002 0.04696 -0.0281 0.0115 1.0000 16.250 1.5525 0.05455 0.05167 -0.0275 0.0115 1.0000 16.500 1.5261 0.06090 0.05823 -0.0288 0.0115 1.0000