XFOIL Version 6.94 Calculated polar for: pp13 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.0113 0.04806 0.03842 -0.0214 0.9998 1.0002 -2.750 0.0115 0.04706 0.03756 -0.0198 0.9998 1.0002 -2.500 0.0112 0.04606 0.03671 -0.0182 0.9998 1.0002 -2.250 0.0090 0.04511 0.03591 -0.0161 0.9998 1.0000 -2.000 -0.0252 0.04421 0.03523 -0.0067 0.9998 0.9884 -1.750 -0.0657 0.04337 0.03458 0.0034 0.9998 0.9741 -1.500 -0.1081 0.04234 0.03374 0.0137 0.9998 0.9618 -1.250 -0.1517 0.04111 0.03269 0.0241 0.9998 0.9520 -1.000 -0.1933 0.03983 0.03154 0.0342 0.9998 0.9422 -0.750 -0.2227 0.03904 0.03077 0.0412 0.9998 0.9233 -0.500 -0.1852 0.04131 0.03272 0.0313 0.9998 0.8601 -0.250 -0.0719 0.04570 0.03666 0.0027 0.9998 0.7447 0.000 0.0158 0.04848 0.03903 -0.0177 0.9998 0.6422 0.250 0.0844 0.05007 0.04025 -0.0318 0.9998 0.5589 0.500 0.1404 0.05097 0.04082 -0.0415 0.9998 0.4974 0.750 0.1884 0.05150 0.04106 -0.0485 0.9998 0.4568 1.000 0.2317 0.05147 0.04085 -0.0536 0.9998 0.4314 1.250 0.2752 0.05171 0.04087 -0.0587 0.9998 0.4088 1.500 0.3184 0.05207 0.04103 -0.0635 0.9998 0.3882 1.750 0.3603 0.05260 0.04136 -0.0679 0.9998 0.3717 2.000 0.4020 0.05294 0.04165 -0.0722 0.9998 0.3638 2.250 0.4436 0.05351 0.04222 -0.0763 0.9998 0.3622 2.500 0.4843 0.05421 0.04300 -0.0801 0.9998 0.3646 2.750 0.5263 0.05497 0.04395 -0.0839 0.9998 0.3673 3.000 0.5721 0.05568 0.04503 -0.0878 0.9998 0.3722 3.250 0.9641 0.04344 0.02970 -0.1174 0.2696 1.0002 3.500 0.9980 0.04520 0.03094 -0.1169 0.2595 1.0002 3.750 1.0377 0.04714 0.03249 -0.1173 0.2529 1.0002 4.000 1.0788 0.04900 0.03444 -0.1179 0.2488 1.0002 4.250 1.1206 0.05116 0.03681 -0.1189 0.2461 1.0002 4.500 1.1593 0.05363 0.03955 -0.1198 0.2431 1.0002 4.750 1.1938 0.05640 0.04262 -0.1202 0.2405 1.0002 5.000 1.2238 0.05946 0.04606 -0.1202 0.2394 1.0002