XFOIL Version 6.96 Calculated polar for: RUTAN CANARD AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -5.000 -0.4905 0.04095 0.03356 -0.0317 0.9915 0.1143 -4.500 -0.4230 0.03431 0.02619 -0.0361 0.9841 0.1094 -4.000 -0.3547 0.03041 0.02147 -0.0394 0.9751 0.1124 -3.500 -0.2853 0.02766 0.01824 -0.0427 0.9670 0.1197 -3.000 -0.2104 0.02595 0.01618 -0.0467 0.9586 0.1315 -2.500 -0.1489 0.02459 0.01483 -0.0484 0.9487 0.1482 -2.000 -0.0797 0.02342 0.01392 -0.0516 0.9400 0.1775 -1.500 -0.0164 0.02239 0.01332 -0.0537 0.9303 0.2310 -1.000 0.0091 0.02033 0.01407 -0.0449 0.9202 0.8747 1.000 0.3060 0.02285 0.01584 -0.0612 0.8775 1.0000 1.500 0.3432 0.02375 0.01666 -0.0592 0.8617 1.0000 2.000 0.3899 0.02469 0.01758 -0.0586 0.8462 1.0000 2.500 0.4582 0.02495 0.01788 -0.0605 0.8279 1.0000 3.000 0.5515 0.02245 0.01545 -0.0615 0.7946 1.0000 3.500 0.6247 0.02009 0.01313 -0.0593 0.7685 1.0000 4.000 0.6752 0.01904 0.01217 -0.0556 0.7373 1.0000 4.500 0.7304 0.01732 0.01051 -0.0515 0.6984 1.0000 5.000 0.7754 0.01595 0.00911 -0.0461 0.6099 1.0000 5.500 0.7808 0.01896 0.00963 -0.0364 0.2438 1.0000 6.000 0.8125 0.02158 0.01160 -0.0328 0.1849 1.0000 6.500 0.8564 0.02395 0.01359 -0.0310 0.1559 1.0000 7.000 0.9070 0.02626 0.01582 -0.0302 0.1368 1.0000 7.500 0.9604 0.02894 0.01844 -0.0298 0.1233 1.0000 8.000 1.0148 0.03218 0.02159 -0.0300 0.1131 1.0000 8.500 1.0612 0.03522 0.02517 -0.0284 0.1063 1.0000 9.000 1.1087 0.03944 0.02947 -0.0280 0.1001 1.0000 9.500 1.1397 0.04405 0.03495 -0.0248 0.0970 1.0000 10.000 1.1633 0.04927 0.04085 -0.0215 0.0935 1.0000