XFOIL Version 6.94 Calculated polar for: nacc632 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0158 0.02482 0.01326 -0.0248 0.9994 1.0006 -2.750 0.0056 0.02533 0.01355 -0.0245 0.9994 1.0006 -2.500 0.0269 0.02591 0.01392 -0.0242 0.9994 1.0006 -2.250 0.0481 0.02654 0.01438 -0.0239 0.9994 1.0006 -2.000 0.0695 0.02722 0.01492 -0.0237 0.9994 1.0006 -1.750 0.0909 0.02795 0.01552 -0.0235 0.9994 1.0006 -1.500 0.1122 0.02872 0.01617 -0.0234 0.9994 1.0006 -1.250 0.1335 0.02953 0.01689 -0.0233 0.9994 1.0006 -1.000 0.1546 0.03038 0.01768 -0.0233 0.9994 1.0006 -0.750 0.1757 0.03128 0.01853 -0.0233 0.9994 1.0006 -0.500 0.1965 0.03222 0.01944 -0.0234 0.9994 1.0006 -0.250 0.2170 0.03322 0.02043 -0.0235 0.9994 1.0006 0.000 0.2373 0.03428 0.02151 -0.0237 0.9994 1.0006 0.250 0.2571 0.03542 0.02268 -0.0239 0.9994 1.0006 0.500 0.2764 0.03666 0.02398 -0.0242 0.9994 1.0006 0.750 0.2948 0.03804 0.02545 -0.0247 0.9994 1.0006 1.000 0.3119 0.03964 0.02719 -0.0253 0.9994 1.0006 1.250 0.4759 0.03709 0.02506 -0.0506 0.9279 1.0006 1.500 0.6054 0.03284 0.02125 -0.0623 0.8304 1.0006 1.750 0.6301 0.03201 0.02035 -0.0553 0.7376 1.0006 2.000 0.6474 0.03180 0.01982 -0.0487 0.6583 1.0006 2.250 0.6682 0.03195 0.01958 -0.0443 0.5986 1.0006 2.500 0.6918 0.03252 0.01979 -0.0417 0.5492 1.0006 2.750 0.7175 0.03347 0.02053 -0.0405 0.5087 1.0006 3.000 0.7444 0.03459 0.02144 -0.0397 0.4781 1.0006 3.250 0.7711 0.03581 0.02247 -0.0390 0.4506 1.0006 3.500 0.7989 0.03733 0.02394 -0.0389 0.4265 1.0006 3.750 0.8266 0.03900 0.02563 -0.0389 0.4061 1.0006 4.000 0.8537 0.04076 0.02735 -0.0388 0.3869 1.0006 4.250 0.8812 0.04288 0.02956 -0.0391 0.3714 1.0006 4.500 0.9085 0.04546 0.03235 -0.0398 0.3598 1.0006 4.750 0.9358 0.04798 0.03492 -0.0401 0.3495 1.0006 5.000 0.9598 0.05098 0.03833 -0.0410 0.3375 1.0006