XFOIL Version 6.8 Calculated polar for: EPPLER 479 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = .000 Re = 3.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top Xtr Bot Xtr ------- -------- --------- --------- -------- ------- ------- -7.000 -.7952 .00815 .00295 -.0016 .6950 .1271 -6.500 -.7407 .00783 .00263 -.0012 .6588 .1454 -6.000 -.6862 .00751 .00232 -.0009 .6227 .1666 -5.500 -.6307 .00725 .00208 -.0006 .5920 .1868 -5.000 -.5747 .00706 .00187 -.0004 .5613 .2040 -4.500 -.5184 .00688 .00169 -.0002 .5316 .2235 -4.000 -.4616 .00674 .00154 -.0001 .5066 .2407 -3.500 -.4050 .00658 .00140 .0001 .4808 .2650 -3.000 -.3475 .00652 .00131 .0001 .4583 .2770 -2.500 -.2900 .00645 .00123 .0001 .4362 .2913 -2.000 -.2321 .00642 .00117 .0001 .4161 .3022 -1.500 -.1740 .00641 .00113 .0001 .3971 .3124 -1.000 -.1162 .00638 .00109 .0001 .3799 .3265 -.500 -.0581 .00638 .00107 .0000 .3644 .3379 .000 .0000 .00640 .00107 .0000 .3496 .3496 .500 .0581 .00638 .00108 .0000 .3378 .3641 1.000 .1162 .00638 .00109 -.0001 .3265 .3798 1.500 .1740 .00641 .00113 -.0001 .3125 .3971 2.000 .2320 .00642 .00117 -.0001 .3024 .4160 2.500 .2899 .00645 .00123 -.0001 .2913 .4361 3.000 .3474 .00652 .00131 -.0001 .2770 .4583 3.500 .4051 .00656 .00140 -.0001 .2666 .4807 4.000 .4624 .00664 .00152 .0000 .2530 .5065 4.500 .5196 .00675 .00165 .0001 .2398 .5318 5.000 .5762 .00689 .00181 .0002 .2235 .5615 5.500 .6324 .00706 .00200 .0004 .2080 .5921 6.000 .6884 .00726 .00222 .0006 .1919 .6230 6.500 .7440 .00747 .00247 .0009 .1768 .6590 7.000 .7996 .00766 .00272 .0011 .1642 .6960 7.500 .8537 .00797 .00305 .0016 .1460 .7401 8.000 .9074 .00827 .00340 .0021 .1305 .7860 8.500 .9598 .00853 .00377 .0029 .1170 .8483 9.000 1.0092 .00873 .00420 .0045 .1045 .9577 9.500 1.0595 .00927 .00470 .0054 .0920 .9788 10.000 1.0989 .01006 .00544 .0085 .0812 .9833