XFOIL Version 6.96 Calculated polar for: AG08 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.500 -0.5792 0.10054 0.09835 0.0184 1.0000 0.0149 -8.250 -0.5757 0.09654 0.09436 0.0162 1.0000 0.0154 -8.000 -0.5725 0.09236 0.09021 0.0133 1.0000 0.0161 -7.750 -0.5681 0.08794 0.08582 0.0091 1.0000 0.0168 -7.500 -0.5579 0.08348 0.08137 0.0037 1.0000 0.0172 -6.000 -0.4718 0.02594 0.02221 -0.0389 1.0000 0.0122 -5.750 -0.4489 0.02142 0.01709 -0.0394 1.0000 0.0122 -5.500 -0.4239 0.02102 0.01669 -0.0395 1.0000 0.0136 -5.250 -0.3984 0.01947 0.01491 -0.0394 1.0000 0.0147 -5.000 -0.3727 0.01625 0.01115 -0.0390 1.0000 0.0148 -4.750 -0.3464 0.01444 0.00901 -0.0386 1.0000 0.0157 -4.500 -0.3198 0.01332 0.00768 -0.0381 1.0000 0.0169 -4.250 -0.2937 0.01184 0.00597 -0.0376 1.0000 0.0183 -4.000 -0.2678 0.01103 0.00513 -0.0372 1.0000 0.0214 -3.750 -0.2414 0.01042 0.00444 -0.0367 1.0000 0.0241 -3.500 -0.2153 0.00974 0.00367 -0.0362 1.0000 0.0271 -3.250 -0.1892 0.00922 0.00313 -0.0358 1.0000 0.0339 -3.000 -0.1631 0.00875 0.00265 -0.0353 1.0000 0.0424 -2.750 -0.1371 0.00841 0.00233 -0.0349 1.0000 0.0549 -2.500 -0.1111 0.00814 0.00211 -0.0345 1.0000 0.0723 -2.250 -0.0851 0.00786 0.00190 -0.0341 1.0000 0.0947 -2.000 -0.0591 0.00761 0.00177 -0.0338 1.0000 0.1267 -1.750 -0.0330 0.00734 0.00167 -0.0335 1.0000 0.1727 -1.500 -0.0042 0.00703 0.00161 -0.0339 0.9990 0.2397 -1.250 0.0356 0.00661 0.00152 -0.0367 0.9936 0.3424 -1.000 0.0738 0.00612 0.00146 -0.0392 0.9861 0.4710 -0.750 0.1059 0.00508 0.00143 -0.0402 0.9760 0.7578 -0.500 0.1442 0.00458 0.00137 -0.0416 0.9633 1.0000 -0.250 0.1778 0.00456 0.00128 -0.0425 0.9419 1.0000 0.000 0.2055 0.00458 0.00120 -0.0421 0.9129 1.0000 0.250 0.2306 0.00465 0.00114 -0.0410 0.8772 1.0000 0.500 0.2555 0.00477 0.00109 -0.0399 0.8384 1.0000 0.750 0.2808 0.00493 0.00106 -0.0390 0.7967 1.0000 1.000 0.3065 0.00512 0.00106 -0.0382 0.7532 1.0000 1.250 0.3325 0.00534 0.00107 -0.0376 0.7077 1.0000 1.500 0.3588 0.00557 0.00110 -0.0371 0.6608 1.0000 1.750 0.3853 0.00582 0.00115 -0.0367 0.6136 1.0000 2.000 0.4118 0.00609 0.00122 -0.0364 0.5659 1.0000 2.250 0.4386 0.00636 0.00130 -0.0361 0.5198 1.0000 2.500 0.4654 0.00663 0.00140 -0.0359 0.4764 1.0000 2.750 0.4923 0.00691 0.00153 -0.0357 0.4359 1.0000 3.000 0.5192 0.00719 0.00166 -0.0355 0.3982 1.0000 3.250 0.5463 0.00745 0.00181 -0.0353 0.3640 1.0000 3.500 0.5733 0.00773 0.00197 -0.0352 0.3312 1.0000 4.000 0.6271 0.00831 0.00235 -0.0349 0.2711 1.0000 4.250 0.6540 0.00861 0.00256 -0.0347 0.2417 1.0000 4.500 0.6808 0.00893 0.00279 -0.0346 0.2131 1.0000 4.750 0.7075 0.00928 0.00307 -0.0344 0.1840 1.0000 5.000 0.7339 0.00967 0.00335 -0.0342 0.1541 1.0000 5.250 0.7602 0.01011 0.00367 -0.0341 0.1240 1.0000 5.500 0.7862 0.01059 0.00403 -0.0339 0.0955 1.0000 5.750 0.8121 0.01112 0.00445 -0.0336 0.0692 1.0000 6.000 0.8376 0.01176 0.00500 -0.0333 0.0457 1.0000 6.250 0.8629 0.01247 0.00565 -0.0330 0.0288 1.0000 6.500 0.8881 0.01325 0.00648 -0.0325 0.0209 1.0000 6.750 0.9113 0.01457 0.00792 -0.0317 0.0165 1.0000 7.000 0.9363 0.01531 0.00878 -0.0312 0.0150 1.0000 7.250 0.9605 0.01619 0.00976 -0.0306 0.0135 1.0000 7.500 0.9843 0.01707 0.01073 -0.0301 0.0120 1.0000 7.750 1.0015 0.01983 0.01375 -0.0287 0.0108 1.0000 8.000 1.0246 0.02098 0.01505 -0.0280 0.0105 1.0000 8.250 1.0467 0.02240 0.01668 -0.0271 0.0100 1.0000 8.500 1.0672 0.02422 0.01873 -0.0262 0.0096 1.0000 8.750 1.0858 0.02650 0.02130 -0.0250 0.0092 1.0000 9.000 1.1014 0.02943 0.02460 -0.0237 0.0091 1.0000 9.250 1.1119 0.03331 0.02893 -0.0222 0.0091 1.0000 9.500 1.1148 0.03826 0.03439 -0.0204 0.0094 1.0000 9.750 1.1096 0.04382 0.04046 -0.0189 0.0097 1.0000 10.000 1.0965 0.04948 0.04647 -0.0178 0.0100 1.0000 10.250 1.0765 0.05448 0.05173 -0.0169 0.0102 1.0000 10.500 1.0553 0.05973 0.05718 -0.0189 0.0103 1.0000 10.750 1.0373 0.06684 0.06446 -0.0254 0.0104 1.0000 11.000 1.0217 0.07624 0.07402 -0.0348 0.0104 1.0000 11.250 1.0038 0.08743 0.08533 -0.0441 0.0103 1.0000