XFOIL Version 6.94 Calculated polar for: pp12 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.026 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.0033 0.04781 0.03910 -0.0184 0.9999 1.0001 -2.750 0.0035 0.04681 0.03822 -0.0168 0.9999 1.0001 -2.500 0.0032 0.04584 0.03740 -0.0151 0.9999 1.0001 -2.250 0.0022 0.04488 0.03657 -0.0135 0.9999 1.0001 -2.000 0.0009 0.04393 0.03577 -0.0117 0.9999 1.0001 -1.750 -0.0012 0.04296 0.03495 -0.0098 0.9999 1.0001 -1.500 -0.0208 0.04195 0.03413 -0.0038 0.9999 0.9956 -1.250 -0.0485 0.04080 0.03315 0.0037 0.9999 0.9824 -1.000 -0.0876 0.03967 0.03219 0.0132 0.9999 0.9710 -0.750 -0.1276 0.03843 0.03108 0.0230 0.9999 0.9618 -0.500 -0.1677 0.03726 0.02992 0.0327 0.9999 0.9478 -0.250 -0.1680 0.03790 0.03037 0.0336 0.9999 0.9123 0.000 -0.0503 0.04394 0.03584 0.0036 0.9999 0.7799 0.250 0.0539 0.04764 0.03913 -0.0215 0.9999 0.6439 0.500 0.1254 0.04978 0.04088 -0.0366 0.9999 0.5412 0.750 0.1798 0.05045 0.04128 -0.0455 0.9999 0.4687 1.000 0.2295 0.05127 0.04172 -0.0528 0.9999 0.4127 1.250 0.2710 0.05134 0.04165 -0.0574 0.9999 0.3855 1.500 0.3129 0.05171 0.04179 -0.0620 0.9999 0.3656 1.750 0.3553 0.05216 0.04206 -0.0663 0.9999 0.3453 2.000 0.3932 0.05277 0.04257 -0.0700 0.9999 0.3316 2.250 0.4361 0.05335 0.04311 -0.0741 0.9999 0.3220 2.500 0.4769 0.05407 0.04388 -0.0778 0.9999 0.3164 2.750 0.8954 0.04104 0.02831 -0.1157 0.2435 0.4527 3.000 0.9691 0.04267 0.02957 -0.1222 0.2306 1.0001 3.250 1.0117 0.04409 0.03034 -0.1225 0.2247 1.0001 3.500 1.0630 0.04577 0.03177 -0.1248 0.2202 1.0001 3.750 1.1305 0.04857 0.03427 -0.1302 0.2114 1.0001 4.000 1.1804 0.05073 0.03666 -0.1328 0.2103 1.0001 4.250 1.2243 0.05346 0.03971 -0.1345 0.2102 1.0001 4.500 1.2609 0.05636 0.04295 -0.1352 0.2120 1.0001 4.750 1.2828 0.05893 0.04617 -0.1334 0.2153 1.0001 5.000 1.3003 0.06172 0.04952 -0.1317 0.2198 1.0001