XFOIL Version 6.96 Calculated polar for: GOE 411 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.000 -0.8809 0.03223 0.02499 0.0012 1.0000 0.0357 -7.750 -0.8562 0.02932 0.02165 0.0030 1.0000 0.0300 -7.500 -0.8398 0.02859 0.02069 0.0057 1.0000 0.0284 -7.250 -0.8231 0.02786 0.01985 0.0079 1.0000 0.0278 -7.000 -0.8024 0.02619 0.01810 0.0093 1.0000 0.0278 -6.750 -0.7828 0.02384 0.01575 0.0106 1.0000 0.0291 -6.500 -0.7701 0.02274 0.01463 0.0134 1.0000 0.0292 -6.250 -0.7602 0.02173 0.01358 0.0167 1.0000 0.0292 -6.000 -0.7520 0.02082 0.01260 0.0202 1.0000 0.0294 -5.750 -0.7443 0.01998 0.01173 0.0238 1.0000 0.0297 -5.500 -0.7363 0.01919 0.01086 0.0273 1.0000 0.0304 -5.250 -0.7256 0.01856 0.01013 0.0304 1.0000 0.0312 -5.000 -0.7135 0.01798 0.00942 0.0332 1.0000 0.0327 -4.750 -0.6999 0.01748 0.00880 0.0357 1.0000 0.0356 -4.500 -0.6849 0.01709 0.00827 0.0380 1.0000 0.0384 -4.000 -0.6340 0.01493 0.00717 0.0373 0.9913 0.2333 -3.750 -0.6086 0.01390 0.00707 0.0368 0.9847 0.3993 -3.500 -0.5793 0.01359 0.00708 0.0362 0.9778 0.4816 -3.250 -0.5450 0.01344 0.00711 0.0346 0.9715 0.5410 -3.000 -0.5152 0.01326 0.00705 0.0341 0.9642 0.5858 -2.750 -0.4804 0.01306 0.00702 0.0327 0.9579 0.6308 -2.500 -0.4482 0.01285 0.00695 0.0318 0.9506 0.6628 -2.250 -0.4098 0.01268 0.00682 0.0297 0.9445 0.6890 -2.000 -0.3747 0.01251 0.00673 0.0283 0.9376 0.7162 -1.750 -0.3361 0.01233 0.00662 0.0262 0.9311 0.7442 -1.500 -0.2958 0.01216 0.00650 0.0238 0.9254 0.7717 -1.250 -0.2574 0.01200 0.00641 0.0219 0.9180 0.7983 -1.000 -0.2071 0.01186 0.00635 0.0177 0.9148 0.8270 -0.750 -0.1635 0.01188 0.00644 0.0151 0.9075 0.8533 -0.500 -0.1106 0.01187 0.00645 0.0104 0.9033 0.8724 -0.250 -0.0492 0.01188 0.00645 0.0039 0.9006 0.8827 0.000 0.0000 0.01192 0.00650 0.0000 0.8912 0.8912 0.250 0.0490 0.01188 0.00645 -0.0039 0.8825 0.9006 0.500 0.1105 0.01187 0.00645 -0.0104 0.8725 0.9033 0.750 0.1635 0.01188 0.00644 -0.0151 0.8531 0.9074 1.000 0.2072 0.01186 0.00635 -0.0178 0.8275 0.9148 1.250 0.2575 0.01200 0.00641 -0.0219 0.7981 0.9180 1.500 0.2959 0.01216 0.00650 -0.0238 0.7716 0.9254 1.750 0.3364 0.01233 0.00662 -0.0262 0.7440 0.9311 2.000 0.3747 0.01251 0.00673 -0.0282 0.7157 0.9376 2.250 0.4099 0.01268 0.00682 -0.0297 0.6888 0.9445 2.500 0.4484 0.01286 0.00695 -0.0318 0.6625 0.9507 2.750 0.4802 0.01306 0.00702 -0.0326 0.6307 0.9579 3.000 0.5149 0.01326 0.00704 -0.0340 0.5845 0.9641 3.250 0.5454 0.01344 0.00711 -0.0347 0.5413 0.9715 3.500 0.5795 0.01358 0.00708 -0.0362 0.4832 0.9777 3.750 0.6085 0.01391 0.00707 -0.0368 0.3972 0.9847 4.000 0.6340 0.01495 0.00718 -0.0373 0.2311 0.9914 4.250 0.6577 0.01661 0.00778 -0.0378 0.0472 0.9973 4.750 0.7000 0.01748 0.00880 -0.0357 0.0358 1.0000 5.000 0.7134 0.01800 0.00944 -0.0331 0.0324 1.0000 5.250 0.7260 0.01854 0.01010 -0.0304 0.0314 1.0000 5.500 0.7354 0.01927 0.01094 -0.0271 0.0301 1.0000 5.750 0.7437 0.02004 0.01178 -0.0237 0.0295 1.0000 6.000 0.7506 0.02095 0.01273 -0.0199 0.0291 1.0000 6.250 0.7600 0.02176 0.01361 -0.0166 0.0292 1.0000 6.500 0.7700 0.02275 0.01463 -0.0134 0.0291 1.0000 6.750 0.7827 0.02404 0.01594 -0.0106 0.0287 1.0000 7.000 0.8004 0.02565 0.01756 -0.0089 0.0283 1.0000 7.250 0.8226 0.02771 0.01971 -0.0078 0.0279 1.0000 7.500 0.8391 0.02840 0.02050 -0.0055 0.0286 1.0000 7.750 0.8566 0.02934 0.02168 -0.0030 0.0301 1.0000 8.000 0.8805 0.03209 0.02484 -0.0012 0.0354 1.0000