XFOIL Version 6.94 Calculated polar for: manu4/38 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.045 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -2.250 -0.0316 0.04297 0.03244 -0.0450 0.4073 0.2376 -2.000 -0.0120 0.04168 0.03075 -0.0435 0.3937 0.2337 -1.750 0.0135 0.04023 0.02887 -0.0430 0.3824 0.2300 -1.500 0.0435 0.03919 0.02730 -0.0433 0.3718 0.2315 -1.250 0.0790 0.03807 0.02575 -0.0446 0.3641 0.2352 -1.000 0.1225 0.03696 0.02405 -0.0474 0.3581 0.2380 -0.750 0.1739 0.03596 0.02267 -0.0519 0.3524 0.2436 -0.500 0.2429 0.03524 0.02137 -0.0600 0.3478 0.2568 -0.250 0.3214 0.03464 0.02059 -0.0704 0.3436 0.2884 0.000 0.4364 0.03295 0.01920 -0.0888 0.3393 0.4369 0.250 0.4960 0.03251 0.01930 -0.0955 0.3375 0.5407 0.500 0.6214 0.03284 0.02045 -0.1171 0.3356 1.0001 0.750 0.6587 0.03373 0.02109 -0.1189 0.3352 1.0001 1.000 0.6909 0.03467 0.02183 -0.1197 0.3346 1.0001 1.250 0.7207 0.03565 0.02268 -0.1201 0.3340 1.0001 1.500 0.7491 0.03670 0.02360 -0.1202 0.3332 1.0001 1.750 0.7766 0.03783 0.02463 -0.1201 0.3324 1.0001 2.000 0.8030 0.03895 0.02572 -0.1198 0.3322 1.0001 2.250 0.8278 0.04001 0.02688 -0.1192 0.3332 1.0001 2.500 0.8515 0.04112 0.02817 -0.1184 0.3358 1.0001 2.750 0.8744 0.04245 0.02964 -0.1176 0.3385 1.0001 3.000 0.8966 0.04393 0.03124 -0.1168 0.3414 1.0001 3.250 0.9181 0.04551 0.03292 -0.1158 0.3445 1.0001 3.500 0.9397 0.04728 0.03478 -0.1150 0.3474 1.0001 3.750 0.9623 0.04932 0.03682 -0.1145 0.3499 1.0001 4.000 0.9747 0.05048 0.03852 -0.1122 0.3580 1.0001 4.250 0.9900 0.05267 0.04092 -0.1107 0.3645 1.0001 4.500 1.0099 0.05509 0.04338 -0.1100 0.3693 1.0001 4.750 1.0133 0.05717 0.04600 -0.1074 0.3828 1.0001 5.000 1.0317 0.06006 0.04898 -0.1069 0.3904 1.0001