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 = 0.026 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.0119 0.04688 0.03809 -0.0209 0.9998 1.0002 -2.750 0.0120 0.04589 0.03723 -0.0193 0.9998 1.0002 -2.500 0.0117 0.04490 0.03638 -0.0176 0.9998 1.0002 -2.250 0.0109 0.04395 0.03556 -0.0159 0.9998 1.0002 -2.000 0.0096 0.04300 0.03475 -0.0141 0.9998 1.0002 -1.750 0.0076 0.04204 0.03394 -0.0122 0.9998 1.0002 -1.500 -0.0101 0.04104 0.03311 -0.0068 0.9998 0.9969 -1.250 -0.0391 0.03985 0.03209 0.0010 0.9998 0.9833 -1.000 -0.0786 0.03871 0.03112 0.0107 0.9998 0.9717 -0.750 -0.1210 0.03740 0.02995 0.0210 0.9998 0.9620 -0.500 -0.1573 0.03626 0.02886 0.0300 0.9998 0.9505 -0.250 -0.1682 0.03653 0.02895 0.0335 0.9998 0.9207 0.000 -0.0663 0.04227 0.03416 0.0076 0.9998 0.7992 0.250 0.0339 0.04639 0.03785 -0.0168 0.9998 0.6712 0.500 0.1070 0.04866 0.03975 -0.0326 0.9998 0.5637 0.750 0.1622 0.04939 0.04021 -0.0420 0.9998 0.4879 1.000 0.2123 0.05005 0.04056 -0.0495 0.9998 0.4280 1.250 0.2570 0.05057 0.04079 -0.0550 0.9998 0.3905 1.500 0.2992 0.05082 0.04084 -0.0596 0.9998 0.3688 1.750 0.3405 0.05129 0.04110 -0.0638 0.9998 0.3493 2.000 0.3838 0.05152 0.04129 -0.0685 0.9998 0.3350 2.250 0.4253 0.05215 0.04181 -0.0724 0.9998 0.3237 2.500 0.4672 0.05277 0.04248 -0.0765 0.9998 0.3174 3.000 0.9383 0.04124 0.02867 -0.1201 0.2346 1.0002 3.250 0.9744 0.04264 0.02915 -0.1193 0.2287 1.0002 3.500 1.0125 0.04407 0.03021 -0.1191 0.2237 1.0002 3.750 1.0560 0.04573 0.03169 -0.1200 0.2195 1.0002 4.000 1.1018 0.04778 0.03363 -0.1216 0.2145 1.0002 4.250 1.1465 0.05022 0.03611 -0.1232 0.2111 1.0002 4.500 1.1829 0.05263 0.03880 -0.1236 0.2105 1.0002 4.750 1.2173 0.05559 0.04199 -0.1238 0.2118 1.0002 5.000 1.2421 0.05799 0.04508 -0.1227 0.2150 1.0002