XFOIL Version 6.94 Calculated polar for: manu4/30 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 -0.0624 0.06165 0.05371 0.0001 0.5001 0.8138 -2.750 -0.1357 0.05974 0.05227 0.0140 0.5097 0.7948 -2.500 -0.1759 0.05792 0.05057 0.0197 0.5107 0.7575 -2.250 -0.1776 0.05742 0.04969 0.0150 0.5030 0.6906 -2.000 -0.1068 0.05878 0.05013 -0.0057 0.4853 0.5656 -1.750 -0.0256 0.05877 0.04905 -0.0245 0.4724 0.4554 -1.500 0.0327 0.05744 0.04704 -0.0338 0.4632 0.4105 -1.250 0.0937 0.05602 0.04485 -0.0422 0.4536 0.3865 -1.000 0.1491 0.05460 0.04287 -0.0491 0.4463 0.3678 -0.750 0.2014 0.05321 0.04107 -0.0545 0.4378 0.3596 -0.500 0.2641 0.05223 0.03944 -0.0620 0.4304 0.3559 -0.250 0.3342 0.05172 0.03822 -0.0709 0.4254 0.3637 0.000 0.3985 0.05089 0.03715 -0.0786 0.4232 0.3798 0.250 0.4725 0.05029 0.03624 -0.0882 0.4215 0.4033 0.500 0.5574 0.04946 0.03534 -0.1001 0.4205 0.4678 0.750 0.6643 0.04635 0.03379 -0.1166 0.4199 0.7198 1.000 0.7681 0.04864 0.03595 -0.1338 0.4208 1.0001 1.250 0.8092 0.05030 0.03738 -0.1370 0.4222 1.0001 1.500 0.8457 0.05205 0.03891 -0.1393 0.4239 1.0001 1.750 0.8794 0.05394 0.04059 -0.1411 0.4257 1.0001 2.000 0.8998 0.05526 0.04199 -0.1402 0.4287 1.0001 2.250 0.9128 0.05679 0.04367 -0.1383 0.4329 1.0001 2.500 0.9264 0.05873 0.04569 -0.1366 0.4369 1.0001 2.750 0.9386 0.06091 0.04792 -0.1349 0.4405 1.0001 3.000 0.9486 0.06330 0.05034 -0.1328 0.4439 1.0001 3.250 0.9589 0.06592 0.05297 -0.1310 0.4472 1.0001 3.500 0.9708 0.06872 0.05573 -0.1296 0.4500 1.0001 3.750 0.9624 0.07156 0.05877 -0.1254 0.4554 1.0001 4.000 0.8969 0.07695 0.06459 -0.1150 0.4674 1.0001 4.250 0.8905 0.08130 0.06892 -0.1127 0.4745 1.0001