XFOIL Version 6.94 Calculated polar for: NLR-7301 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.2694 0.01352 0.00869 -0.0675 0.8004 0.6802 0.500 0.3268 0.01293 0.00805 -0.0668 0.7753 0.6853 1.000 0.3798 0.01255 0.00762 -0.0650 0.7285 0.6892 1.500 0.4210 0.01301 0.00725 -0.0610 0.5357 0.6941 2.000 0.4428 0.01546 0.00808 -0.0559 0.2224 0.6996 2.500 0.4909 0.01646 0.00854 -0.0552 0.1467 0.7056 3.000 0.5406 0.01706 0.00895 -0.0543 0.1237 0.7101 3.500 0.5887 0.01772 0.00954 -0.0528 0.1087 0.7142 4.000 0.6370 0.01847 0.01021 -0.0514 0.0967 0.7190 4.500 0.6856 0.01933 0.01097 -0.0503 0.0879 0.7250 5.000 0.7351 0.02036 0.01188 -0.0496 0.0820 0.7311 5.500 0.7834 0.02111 0.01263 -0.0483 0.0776 0.7359 6.000 0.8302 0.02240 0.01386 -0.0468 0.0745 0.7408 6.500 0.8803 0.02343 0.01496 -0.0458 0.0721 0.7467 7.000 0.9319 0.02459 0.01611 -0.0453 0.0699 0.7536 7.500 0.9830 0.02595 0.01740 -0.0448 0.0679 0.7594 8.000 1.0328 0.02755 0.01913 -0.0438 0.0665 0.7649 8.500 1.0821 0.02919 0.02094 -0.0429 0.0652 0.7715 9.000 1.1318 0.03108 0.02296 -0.0424 0.0641 0.7782 9.500 1.1761 0.03299 0.02505 -0.0409 0.0632 0.7840 10.000 1.2172 0.03516 0.02741 -0.0391 0.0624 0.7908 10.500 1.2568 0.03765 0.03006 -0.0375 0.0618 0.7988 11.000 1.2907 0.04051 0.03314 -0.0353 0.0613 0.8052 11.500 1.3163 0.04423 0.03714 -0.0323 0.0608 0.8120 12.000 1.3227 0.04897 0.04227 -0.0277 0.0604 0.8195 12.500 1.3066 0.05343 0.04723 -0.0213 0.0602 0.8263 13.000 1.2806 0.05952 0.05381 -0.0161 0.0602 0.8334 13.500 1.2442 0.06746 0.06219 -0.0130 0.0603 0.8406