XFOIL Version 6.96 Calculated polar for: NLR-7301 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 ------ -------- --------- --------- -------- -------- -------- -19.750 -1.2614 0.08570 0.08157 -0.0471 1.0000 0.0289 -19.500 -1.2661 0.08152 0.07732 -0.0490 1.0000 0.0290 -19.250 -1.2695 0.07761 0.07334 -0.0507 1.0000 0.0292 -19.000 -1.2717 0.07395 0.06960 -0.0522 1.0000 0.0293 -18.750 -1.2730 0.07043 0.06601 -0.0537 1.0000 0.0295 -18.500 -1.2734 0.06707 0.06258 -0.0550 1.0000 0.0297 -18.250 -1.2727 0.06395 0.05938 -0.0561 1.0000 0.0299 -18.000 -1.2700 0.06112 0.05648 -0.0571 1.0000 0.0301 -17.750 -1.2669 0.05836 0.05364 -0.0580 1.0000 0.0304 -17.500 -1.2626 0.05579 0.05101 -0.0587 1.0000 0.0306 -17.250 -1.2570 0.05343 0.04858 -0.0594 1.0000 0.0308 -17.000 -1.2511 0.05114 0.04622 -0.0600 1.0000 0.0309 -16.750 -1.2432 0.04909 0.04410 -0.0604 1.0000 0.0311 -16.500 -1.2409 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-0.0412 0.0431 0.7480 12.250 1.4295 0.02726 0.02075 -0.0398 0.0430 0.7510 12.500 1.4444 0.02813 0.02166 -0.0385 0.0429 0.7537 12.750 1.4584 0.02907 0.02265 -0.0372 0.0427 0.7558 13.000 1.4722 0.03002 0.02366 -0.0359 0.0426 0.7593 13.250 1.4856 0.03099 0.02471 -0.0346 0.0424 0.7626 13.500 1.4982 0.03204 0.02582 -0.0333 0.0422 0.7657 13.750 1.5107 0.03312 0.02696 -0.0320 0.0420 0.7687 14.000 1.5219 0.03430 0.02819 -0.0308 0.0419 0.7717 14.250 1.5326 0.03554 0.02949 -0.0295 0.0417 0.7744 14.500 1.5428 0.03686 0.03086 -0.0283 0.0416 0.7768 14.750 1.5522 0.03823 0.03232 -0.0271 0.0415 0.7808 15.000 1.5609 0.03967 0.03383 -0.0260 0.0414 0.7843 15.250 1.5685 0.04124 0.03547 -0.0249 0.0413 0.7878 15.500 1.5760 0.04285 0.03715 -0.0238 0.0412 0.7911 15.750 1.5822 0.04459 0.03896 -0.0228 0.0410 0.7941