XFOIL Version 6.96 Calculated polar for: RUTAN WING AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -5.000 -0.4150 0.01217 0.00538 -0.0349 0.9014 0.0640 -4.500 -0.3627 0.01161 0.00480 -0.0341 0.8908 0.0712 -4.000 -0.3083 0.01110 0.00432 -0.0339 0.8795 0.0813 -3.500 -0.2537 0.01065 0.00391 -0.0336 0.8694 0.0944 -3.000 -0.1982 0.01024 0.00355 -0.0335 0.8595 0.1124 -2.500 -0.1423 0.00980 0.00324 -0.0336 0.8498 0.1440 -2.000 -0.0876 0.00901 0.00291 -0.0337 0.8407 0.2523 -1.000 0.0187 0.00716 0.00274 -0.0330 0.8222 0.7199 -0.500 0.0752 0.00714 0.00281 -0.0326 0.8101 0.7628 0.000 0.1305 0.00709 0.00278 -0.0317 0.7921 0.7920 0.500 0.1857 0.00705 0.00272 -0.0308 0.7705 0.8170 1.000 0.2409 0.00706 0.00275 -0.0300 0.7513 0.8411 1.500 0.2944 0.00708 0.00279 -0.0287 0.7265 0.8634 2.000 0.3502 0.00710 0.00273 -0.0282 0.6853 0.8766 2.500 0.4032 0.00728 0.00273 -0.0271 0.6150 0.8904 3.000 0.4426 0.00920 0.00326 -0.0246 0.2871 0.9053 3.500 0.4879 0.01038 0.00383 -0.0229 0.1506 0.9210 4.000 0.5355 0.01098 0.00424 -0.0213 0.1136 0.9393 4.500 0.5843 0.01142 0.00464 -0.0197 0.0955 0.9599 5.000 0.6435 0.01201 0.00517 -0.0207 0.0827 0.9817 5.500 0.7096 0.01273 0.00582 -0.0234 0.0727 1.0000 6.000 0.7604 0.01351 0.00656 -0.0231 0.0659 1.0000 6.500 0.8107 0.01447 0.00746 -0.0227 0.0603 1.0000 7.000 0.8626 0.01517 0.00816 -0.0223 0.0560 1.0000 7.500 0.9107 0.01626 0.00924 -0.0214 0.0523 1.0000 8.000 0.9604 0.01707 0.01006 -0.0207 0.0489 1.0000 8.500 1.0061 0.01836 0.01136 -0.0195 0.0461 1.0000 9.000 1.0534 0.01931 0.01238 -0.0184 0.0434 1.0000 9.500 1.0966 0.02080 0.01386 -0.0170 0.0410 1.0000 10.000 1.1408 0.02195 0.01517 -0.0156 0.0390 1.0000