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 = 0.500 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -19.750 -1.1324 0.10474 0.10012 -0.0385 1.0000 0.0286 -19.500 -1.1423 0.09935 0.09463 -0.0414 1.0000 0.0288 -19.250 -1.1511 0.09423 0.08940 -0.0440 1.0000 0.0289 -19.000 -1.1587 0.08939 0.08446 -0.0464 1.0000 0.0291 -18.750 -1.1649 0.08482 0.07979 -0.0486 1.0000 0.0293 -18.500 -1.1698 0.08051 0.07539 -0.0507 1.0000 0.0295 -18.250 -1.1735 0.07647 0.07125 -0.0525 1.0000 0.0296 -18.000 -1.1757 0.07272 0.06740 -0.0541 1.0000 0.0298 -17.750 -1.1765 0.06921 0.06380 -0.0556 1.0000 0.0300 -17.500 -1.1762 0.06594 0.06044 -0.0568 1.0000 0.0303 -17.250 -1.1749 0.06286 0.05727 -0.0579 1.0000 0.0305 -17.000 -1.1726 0.05997 0.05429 -0.0589 1.0000 0.0307 -16.750 -1.1685 0.05734 0.05157 -0.0597 1.0000 0.0309 -16.500 -1.1639 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-0.0349 0.0419 0.7433 12.500 1.3918 0.03224 0.02480 -0.0336 0.0417 0.7463 12.750 1.4046 0.03332 0.02594 -0.0324 0.0416 0.7493 13.000 1.4171 0.03444 0.02712 -0.0312 0.0414 0.7525 13.250 1.4285 0.03566 0.02839 -0.0300 0.0412 0.7555 13.500 1.4395 0.03690 0.02971 -0.0288 0.0411 0.7584 13.750 1.4497 0.03823 0.03112 -0.0277 0.0409 0.7615 14.000 1.4594 0.03963 0.03259 -0.0266 0.0408 0.7647 14.250 1.4680 0.04112 0.03417 -0.0255 0.0407 0.7678 14.500 1.4762 0.04270 0.03581 -0.0245 0.0405 0.7709 14.750 1.4833 0.04440 0.03758 -0.0235 0.0404 0.7740 15.000 1.4899 0.04618 0.03943 -0.0227 0.0403 0.7770 15.250 1.4949 0.04810 0.04145 -0.0218 0.0402 0.7801 15.500 1.4996 0.05013 0.04356 -0.0211 0.0401 0.7835 15.750 1.5034 0.05231 0.04583 -0.0205 0.0400 0.7869 16.000 1.5063 0.05466 0.04826 -0.0200 0.0399 0.7904 16.250 1.5085 0.05715 0.05083 -0.0197 0.0398 0.7938 16.500 1.5096 0.05982 0.05359 -0.0195 0.0396 0.7971 16.750 1.5094 0.06272 0.05659 -0.0194 0.0395 0.8004 17.000 1.5087 0.06582 0.05981 -0.0197 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