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 = 1.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -5.000 -0.3403 0.01005 0.00393 -0.0469 0.8129 0.0408 -4.500 -0.2839 0.00960 0.00342 -0.0470 0.8035 0.0438 -4.000 -0.2263 0.00925 0.00308 -0.0473 0.7949 0.0472 -3.500 -0.1691 0.00891 0.00272 -0.0475 0.7860 0.0526 -3.000 -0.1111 0.00859 0.00245 -0.0478 0.7775 0.0604 -2.500 -0.0534 0.00834 0.00220 -0.0481 0.7686 0.0713 -2.000 0.0050 0.00807 0.00202 -0.0486 0.7601 0.0853 -1.000 0.1217 0.00761 0.00175 -0.0494 0.7418 0.1352 -0.500 0.1797 0.00728 0.00165 -0.0499 0.7329 0.2134 0.500 0.2916 0.00545 0.00155 -0.0506 0.7023 0.7737 1.000 0.3490 0.00545 0.00158 -0.0505 0.6821 0.8123 1.500 0.4063 0.00549 0.00163 -0.0505 0.6586 0.8412 2.000 0.4625 0.00556 0.00171 -0.0502 0.6316 0.8690 2.500 0.5176 0.00573 0.00181 -0.0497 0.5919 0.8949 3.000 0.5692 0.00613 0.00199 -0.0486 0.5145 0.9187 3.500 0.6131 0.00763 0.00258 -0.0469 0.2990 0.9382 4.000 0.6570 0.00874 0.00313 -0.0449 0.1684 0.9634 4.500 0.7155 0.00936 0.00351 -0.0458 0.1195 0.9909 5.000 0.7737 0.00992 0.00390 -0.0468 0.0943 1.0000 5.500 0.8280 0.01049 0.00435 -0.0468 0.0782 1.0000 6.000 0.8819 0.01102 0.00481 -0.0468 0.0675 1.0000 6.500 0.9346 0.01161 0.00533 -0.0465 0.0586 1.0000 7.000 0.9862 0.01225 0.00593 -0.0461 0.0522 1.0000 7.500 1.0363 0.01298 0.00661 -0.0454 0.0470 1.0000 8.000 1.0869 0.01358 0.00723 -0.0448 0.0441 1.0000 8.500 1.1334 0.01449 0.00813 -0.0436 0.0409 1.0000 9.000 1.1816 0.01516 0.00884 -0.0426 0.0390 1.0000 9.500 1.2264 0.01604 0.00972 -0.0412 0.0370 1.0000 10.000 1.2675 0.01710 0.01083 -0.0393 0.0352 1.0000