XFOIL Version 6.96 Calculated polar for: NACA 65-209 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -7.750 -0.5821 0.04972 0.04770 -0.0419 1.0000 0.0110 -7.500 -0.5841 0.04568 0.04350 -0.0398 1.0000 0.0110 -7.250 -0.5701 0.03953 0.03704 -0.0414 0.9956 0.0111 -7.000 -0.5636 0.02821 0.02502 -0.0441 0.9868 0.0092 -6.750 -0.5406 0.01983 0.01580 -0.0447 0.9818 0.0092 -6.500 -0.5126 0.01761 0.01329 -0.0453 0.9753 0.0099 -6.250 -0.4819 0.01588 0.01129 -0.0462 0.9701 0.0107 -5.750 -0.4307 0.01253 0.00750 -0.0459 0.9528 0.0121 -5.500 -0.4056 0.01193 0.00684 -0.0455 0.9434 0.0127 -5.250 -0.3804 0.01150 0.00636 -0.0451 0.9341 0.0135 -5.000 -0.3550 0.01113 0.00593 -0.0447 0.9254 0.0146 -4.750 -0.3294 0.01082 0.00554 -0.0443 0.9163 0.0156 -4.500 -0.3035 0.01048 0.00514 -0.0439 0.9077 0.0162 -4.000 -0.2551 0.00887 0.00332 -0.0428 0.8905 0.0184 -3.750 -0.2287 0.00862 0.00302 -0.0426 0.8825 0.0203 -3.500 -0.2020 0.00837 0.00272 -0.0425 0.8742 0.0221 -3.250 -0.1751 0.00814 0.00245 -0.0424 0.8664 0.0235 -3.000 -0.1480 0.00796 0.00221 -0.0422 0.8587 0.0247 -2.750 -0.1212 0.00758 0.00176 -0.0421 0.8508 0.0308 -2.250 -0.0669 0.00710 0.00135 -0.0421 0.8354 0.0774 -2.000 -0.0431 0.00588 0.00104 -0.0422 0.8281 0.3568 -1.750 -0.0182 0.00504 0.00090 -0.0422 0.8202 0.5738 -1.500 0.0089 0.00487 0.00087 -0.0422 0.8132 0.6342 -1.250 0.0364 0.00478 0.00086 -0.0421 0.8056 0.6728 -1.000 0.0637 0.00471 0.00088 -0.0420 0.7987 0.7112 -0.750 0.0914 0.00470 0.00089 -0.0420 0.7912 0.7341 -0.500 0.1191 0.00469 0.00090 -0.0419 0.7843 0.7504 -0.250 0.1469 0.00469 0.00091 -0.0419 0.7769 0.7643 0.000 0.1750 0.00470 0.00092 -0.0420 0.7701 0.7726 0.250 0.2031 0.00471 0.00093 -0.0421 0.7627 0.7800 0.500 0.2312 0.00473 0.00094 -0.0422 0.7555 0.7880 0.750 0.2590 0.00474 0.00096 -0.0422 0.7470 0.7953 1.000 0.2869 0.00477 0.00098 -0.0422 0.7346 0.8034 1.250 0.3140 0.00480 0.00100 -0.0420 0.7184 0.8108 1.500 0.3414 0.00486 0.00102 -0.0419 0.6983 0.8188 1.750 0.3681 0.00493 0.00105 -0.0417 0.6728 0.8267 2.000 0.3947 0.00505 0.00109 -0.0414 0.6384 0.8347 2.250 0.4209 0.00522 0.00117 -0.0411 0.5988 0.8431 2.500 0.4452 0.00558 0.00128 -0.0406 0.5200 0.8512 2.750 0.4640 0.00677 0.00165 -0.0394 0.3121 0.8607 3.250 0.5060 0.00859 0.00243 -0.0377 0.0389 0.8793 3.500 0.5314 0.00893 0.00279 -0.0372 0.0283 0.8889 3.750 0.5570 0.00909 0.00300 -0.0368 0.0255 0.8986 4.000 0.5816 0.00938 0.00332 -0.0362 0.0217 0.9086 4.250 0.6043 0.00995 0.00402 -0.0352 0.0192 0.9194 4.500 0.6286 0.01019 0.00432 -0.0345 0.0186 0.9305 4.750 0.6513 0.01049 0.00468 -0.0335 0.0177 0.9430 5.000 0.6729 0.01081 0.00509 -0.0322 0.0168 0.9584 5.250 0.6998 0.01114 0.00546 -0.0322 0.0156 0.9769 5.500 0.7303 0.01162 0.00596 -0.0332 0.0144 1.0000 5.750 0.7528 0.01263 0.00704 -0.0325 0.0136 1.0000 6.000 0.7725 0.01447 0.00900 -0.0313 0.0129 1.0000 6.250 0.7983 0.01498 0.00957 -0.0311 0.0126 1.0000 6.500 0.8240 0.01550 0.01014 -0.0309 0.0121 1.0000 6.750 0.8496 0.01594 0.01064 -0.0307 0.0113 1.0000 7.000 0.8742 0.01671 0.01149 -0.0304 0.0106 1.0000 7.250 0.8984 0.01752 0.01238 -0.0300 0.0101 1.0000 7.500 0.9224 0.01822 0.01314 -0.0297 0.0097 1.0000 7.750 0.9455 0.01918 0.01418 -0.0292 0.0092 1.0000 8.250 0.9800 0.02504 0.02071 -0.0266 0.0083 1.0000 8.500 1.0010 0.02625 0.02210 -0.0257 0.0078 1.0000 8.750 1.0163 0.02897 0.02515 -0.0242 0.0072 1.0000 9.000 1.0075 0.03764 0.03454 -0.0197 0.0069 1.0000 9.250 0.9888 0.04678 0.04424 -0.0151 0.0070 1.0000 9.500 0.9718 0.05345 0.05124 -0.0117 0.0073 1.0000 9.750 0.9550 0.05859 0.05659 -0.0090 0.0073 1.0000 10.000 0.9304 0.06224 0.06038 -0.0055 0.0076 1.0000 10.250 0.9061 0.06618 0.06444 -0.0043 0.0078 1.0000 10.500 0.8843 0.07130 0.06966 -0.0052 0.0078 1.0000 10.750 0.8581 0.07760 0.07606 -0.0086 0.0080 1.0000