XFOIL Version 6.96 Calculated polar for: FX 63-143 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 ------ -------- --------- --------- -------- -------- -------- -10.000 -0.3811 0.08955 0.08745 -0.0414 1.0000 0.0179 -9.750 -0.3879 0.08382 0.08176 -0.0445 1.0000 0.0181 -9.500 -0.3998 0.07641 0.07439 -0.0495 1.0000 0.0182 -9.250 -0.4167 0.06885 0.06683 -0.0560 1.0000 0.0181 -9.000 -0.4328 0.06218 0.06007 -0.0614 0.9955 0.0181 -8.750 -0.4557 0.05194 0.04929 -0.0680 0.9366 0.0184 -8.500 -0.4169 0.04276 0.03944 -0.0811 0.8957 0.0186 -8.250 -0.3950 0.03972 0.03604 -0.0833 0.8364 0.0187 -8.000 -0.3900 0.03760 0.03363 -0.0810 0.8007 0.0188 -7.750 -0.3835 0.03568 0.03144 -0.0787 0.7735 0.0189 -7.500 -0.3759 0.03354 0.02902 -0.0763 0.7519 0.0191 -7.250 -0.3666 0.03133 0.02651 -0.0740 0.7343 0.0193 -7.000 -0.3550 0.02909 0.02396 -0.0718 0.7202 0.0196 -6.750 -0.3411 0.02702 0.02156 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3.500 0.6866 0.00969 0.00309 -0.0509 0.3333 0.8466 3.750 0.7067 0.00999 0.00338 -0.0492 0.3164 0.8717 4.000 0.7275 0.01043 0.00374 -0.0477 0.2956 0.8931 4.250 0.7425 0.01072 0.00403 -0.0448 0.2873 0.9035 4.500 0.7644 0.01097 0.00426 -0.0436 0.2805 0.9117 4.750 0.7790 0.01116 0.00448 -0.0406 0.2756 0.9192 5.000 0.8026 0.01136 0.00469 -0.0399 0.2661 0.9265 5.250 0.8153 0.01154 0.00486 -0.0366 0.2489 0.9319 5.500 0.8316 0.01180 0.00508 -0.0344 0.2390 0.9388 5.750 0.8401 0.01191 0.00519 -0.0302 0.2334 0.9482 6.000 0.8531 0.01211 0.00538 -0.0272 0.2235 0.9571 6.250 0.8651 0.01219 0.00547 -0.0240 0.2148 0.9636 6.500 0.8768 0.01262 0.00577 -0.0213 0.1831 0.9698 6.750 0.8956 0.01285 0.00598 -0.0199 0.1736 0.9739 7.000 0.9168 0.01309 0.00623 -0.0190 0.1654 0.9783 7.250 0.9341 0.01349 0.00657 -0.0176 0.1386 0.9827 7.500 0.9672 0.01405 0.00707 -0.0195 0.1291 0.9887 7.750 1.0105 0.01446 0.00752 -0.0233 0.1231 0.9935 8.250 1.0721 0.01569 0.00867 -0.0267 0.0923 1.0000 8.500 1.0640 0.01590 0.00890 -0.0203 0.0903 1.0000 8.750 1.0691 0.01622 0.00927 -0.0167 0.0881 1.0000 9.000 1.0847 0.01659 0.00971 -0.0152 0.0864 1.0000 9.250 1.1031 0.01699 0.01019 -0.0143 0.0841 1.0000 9.500 1.1196 0.01757 0.01083 -0.0134 0.0797 1.0000 9.750 1.1389 0.01805 0.01139 -0.0129 0.0744 1.0000 10.000 1.1531 0.01886 0.01218 -0.0120 0.0643 1.0000 10.250 1.1632 0.01998 0.01329 -0.0109 0.0597 1.0000 10.500 1.1709 0.02137 0.01472 -0.0099 0.0567 1.0000 10.750 1.1807 0.02271 0.01615 -0.0093 0.0532 1.0000 11.000 1.1897 0.02422 0.01778 -0.0089 0.0493 1.0000 11.250 1.2073 0.02511 0.01874 -0.0090 0.0367 1.0000 11.500 1.2142 0.02692 0.02046 -0.0087 0.0295 1.0000 11.750 1.2198 0.02892 0.02251 -0.0085 0.0271 1.0000 12.000 1.2240 0.03110 0.02475 -0.0084 0.0253 1.0000 12.250 1.2267 0.03349 0.02724 -0.0083 0.0238 1.0000 12.500 1.2267 0.03618 0.03004 -0.0082 0.0228 1.0000 12.750 1.2327 0.03833 0.03227 -0.0084 0.0223 1.0000 13.000 1.2381 0.04054 0.03458 -0.0086 0.0215 1.0000 13.250 1.2412 0.04302 0.03716 -0.0088 0.0211 1.0000 13.500 1.2442 0.04554 0.03977 -0.0091 0.0203 1.0000 13.750 1.2462 0.04823 0.04255 -0.0094 0.0198 1.0000 14.000 1.2476 0.05104 0.04545 -0.0099 0.0192 1.0000 14.250 1.2487 0.05397 0.04847 -0.0104 0.0186 1.0000 14.500 1.2484 0.05713 0.05171 -0.0111 0.0182 1.0000 14.750 1.2468 0.06051 0.05518 -0.0118 0.0176 1.0000 15.000 1.2435 0.06422 0.05897 -0.0127 0.0171 1.0000 15.250 1.2369 0.06836 0.06323 -0.0137 0.0165 1.0000 15.500 1.2347 0.07203 0.06699 -0.0147 0.0162 1.0000 15.750 1.2389 0.07500 0.07006 -0.0157 0.0159 1.0000 16.000 1.2420 0.07813 0.07329 -0.0168 0.0154 1.0000 16.250 1.2444 0.08142 0.07668 -0.0180 0.0148 1.0000 16.500 1.2457 0.08490 0.08025 -0.0193 0.0141 1.0000 16.750 1.2461 0.08858 0.08401 -0.0207 0.0133 1.0000 17.000 1.2439 0.09268 0.08818 -0.0223 0.0127 1.0000 17.250 1.2363 0.09749 0.09306 -0.0240 0.0118 1.0000