XFOIL Version 6.96 Calculated polar for: NASA/AMES A-03 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -10.250 -0.6928 0.10112 0.09800 0.0197 0.9027 0.0160 -10.000 -0.7005 0.09432 0.09122 0.0155 0.9001 0.0160 -9.750 -0.7169 0.08363 0.08054 0.0075 0.8974 0.0159 -9.250 -0.7545 0.06596 0.06259 -0.0065 0.8904 0.0156 -9.000 -0.7776 0.05710 0.05334 -0.0085 0.8870 0.0157 -8.750 -0.7944 0.04895 0.04460 -0.0085 0.8841 0.0159 -8.500 -0.7893 0.04550 0.04087 -0.0078 0.8818 0.0162 -8.250 -0.7773 0.04210 0.03716 -0.0078 0.8791 0.0165 -8.000 -0.7639 0.03838 0.03302 -0.0075 0.8765 0.0168 -7.750 -0.7471 0.03532 0.02957 -0.0071 0.8741 0.0171 -7.500 -0.7280 0.03265 0.02651 -0.0066 0.8721 0.0175 -7.250 -0.7074 0.03020 0.02365 -0.0061 0.8702 0.0180 -7.000 -0.6855 0.02804 0.02109 -0.0054 0.8685 0.0187 -6.750 -0.6620 0.02619 0.01887 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3.000 0.4113 0.01185 0.00376 -0.0021 0.2733 1.0000 3.250 0.4388 0.01212 0.00394 -0.0024 0.2612 1.0000 3.500 0.4667 0.01240 0.00417 -0.0028 0.2522 1.0000 3.750 0.4946 0.01268 0.00439 -0.0032 0.2434 1.0000 4.000 0.5226 0.01296 0.00463 -0.0035 0.2358 1.0000 4.250 0.5505 0.01324 0.00490 -0.0039 0.2288 1.0000 4.500 0.5784 0.01355 0.00517 -0.0042 0.2224 1.0000 4.750 0.6064 0.01382 0.00545 -0.0045 0.2157 1.0000 5.000 0.6340 0.01419 0.00577 -0.0048 0.2096 1.0000 5.250 0.6620 0.01443 0.00607 -0.0051 0.2032 1.0000 5.500 0.6896 0.01476 0.00638 -0.0053 0.1971 1.0000 5.750 0.7172 0.01507 0.00672 -0.0056 0.1910 1.0000 6.000 0.7448 0.01535 0.00706 -0.0058 0.1842 1.0000 6.250 0.7721 0.01570 0.00741 -0.0061 0.1781 1.0000 6.500 0.7996 0.01597 0.00777 -0.0063 0.1711 1.0000 6.750 0.8266 0.01633 0.00813 -0.0065 0.1646 1.0000 7.000 0.8539 0.01661 0.00852 -0.0067 0.1570 1.0000 7.250 0.8807 0.01697 0.00889 -0.0069 0.1498 1.0000 7.500 0.9075 0.01729 0.00927 -0.0071 0.1405 1.0000 7.750 0.9340 0.01764 0.00970 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12.750 1.2622 0.04082 0.03439 0.0039 0.0196 1.0000 13.000 1.2618 0.04338 0.03711 0.0046 0.0191 1.0000 13.250 1.2615 0.04606 0.03997 0.0051 0.0187 1.0000 13.500 1.2591 0.04907 0.04318 0.0052 0.0183 1.0000 13.750 1.2547 0.05249 0.04678 0.0049 0.0178 1.0000 14.000 1.2482 0.05639 0.05086 0.0041 0.0175 1.0000 14.250 1.2396 0.06090 0.05555 0.0025 0.0171 1.0000 14.500 1.2287 0.06620 0.06106 0.0001 0.0169 1.0000 14.750 1.2159 0.07242 0.06747 -0.0034 0.0167 1.0000 15.000 1.2014 0.07955 0.07478 -0.0077 0.0166 1.0000 15.250 1.1853 0.08726 0.08266 -0.0123 0.0165 1.0000 15.500 1.1679 0.09526 0.09083 -0.0169 0.0165 1.0000 15.750 1.1496 0.10337 0.09907 -0.0214 0.0164 1.0000 16.000 1.1308 0.11155 0.10738 -0.0259 0.0164 1.0000 16.250 1.1119 0.11988 0.11584 -0.0304 0.0164 1.0000 16.500 1.0923 0.12848 0.12458 -0.0352 0.0164 1.0000 16.750 1.0711 0.13794 0.13417 -0.0405 0.0164 1.0000 17.000 1.0442 0.14948 0.14586 -0.0471 0.0166 1.0000