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 = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -5.000 -0.3235 0.01307 0.00623 -0.0511 0.8567 0.0504 -4.500 -0.2679 0.01250 0.00559 -0.0510 0.8465 0.0538 -4.000 -0.2136 0.01179 0.00489 -0.0507 0.8374 0.0585 -3.500 -0.1579 0.01120 0.00432 -0.0507 0.8275 0.0649 -3.000 -0.1019 0.01075 0.00387 -0.0507 0.8190 0.0740 -2.500 -0.0450 0.01035 0.00351 -0.0508 0.8096 0.0872 -2.000 0.0118 0.01000 0.00321 -0.0510 0.8009 0.1047 -1.000 0.1260 0.00915 0.00277 -0.0516 0.7826 0.2089 0.000 0.2307 0.00718 0.00283 -0.0503 0.7642 0.8103 0.500 0.2856 0.00719 0.00291 -0.0495 0.7546 0.8506 1.000 0.3385 0.00712 0.00284 -0.0480 0.7357 0.8814 1.500 0.3898 0.00708 0.00279 -0.0462 0.7164 0.9111 2.000 0.4374 0.00702 0.00274 -0.0435 0.6957 0.9408 2.500 0.4890 0.00695 0.00268 -0.0418 0.6703 0.9714 3.000 0.5575 0.00702 0.00268 -0.0443 0.6382 0.9850 3.500 0.6291 0.00730 0.00273 -0.0478 0.5759 0.9958 4.000 0.6757 0.00855 0.00313 -0.0470 0.3907 1.0000 4.500 0.7140 0.01064 0.00413 -0.0452 0.1803 1.0000 5.000 0.7635 0.01162 0.00477 -0.0447 0.1250 1.0000 5.500 0.8139 0.01245 0.00543 -0.0441 0.1000 1.0000 6.000 0.8645 0.01320 0.00610 -0.0435 0.0846 1.0000 6.500 0.9136 0.01403 0.00687 -0.0427 0.0735 1.0000 7.000 0.9611 0.01494 0.00774 -0.0416 0.0653 1.0000 7.500 1.0062 0.01600 0.00877 -0.0402 0.0596 1.0000 8.000 1.0517 0.01692 0.00969 -0.0389 0.0552 1.0000 8.500 1.0931 0.01812 0.01093 -0.0370 0.0517 1.0000 9.000 1.1346 0.01920 0.01204 -0.0351 0.0489 1.0000 9.500 1.1687 0.02073 0.01359 -0.0322 0.0463 1.0000 10.000 1.2041 0.02193 0.01491 -0.0295 0.0442 1.0000