XFOIL Version 6.96 Calculated polar for: NASA/LANGLEY 64-012 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 ------ -------- --------- --------- -------- -------- -------- -12.750 -0.7959 0.08606 0.08235 -0.0117 1.0000 0.0184 -12.500 -0.8464 0.07280 0.06884 -0.0207 1.0000 0.0179 -12.250 -0.8798 0.06477 0.06056 -0.0250 1.0000 0.0177 -12.000 -0.9028 0.05918 0.05474 -0.0270 1.0000 0.0177 -11.750 -0.9208 0.05474 0.05009 -0.0277 1.0000 0.0178 -11.500 -0.9311 0.05158 0.04675 -0.0274 1.0000 0.0180 -11.250 -0.9433 0.04830 0.04323 -0.0264 1.0000 0.0182 -11.000 -0.9522 0.04544 0.04013 -0.0246 1.0000 0.0183 -10.750 -0.9586 0.04280 0.03724 -0.0222 1.0000 0.0185 -10.500 -0.9587 0.04017 0.03433 -0.0202 1.0000 0.0188 -10.250 -0.9535 0.03763 0.03148 -0.0186 1.0000 0.0191 -10.000 -0.9440 0.03532 0.02885 -0.0172 1.0000 0.0195 -9.750 -0.9311 0.03320 0.02642 -0.0160 1.0000 0.0200 -9.500 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0.500 0.0557 0.01014 0.00324 0.0001 0.7039 0.7313 0.750 0.0837 0.01015 0.00327 0.0000 0.6972 0.7387 1.250 0.1394 0.01020 0.00339 0.0001 0.6844 0.7536 1.500 0.1669 0.01023 0.00346 0.0003 0.6783 0.7610 1.750 0.1948 0.01027 0.00355 0.0002 0.6721 0.7692 2.000 0.2223 0.01031 0.00365 0.0004 0.6647 0.7765 2.250 0.2499 0.01035 0.00374 0.0005 0.6567 0.7847 2.500 0.2771 0.01039 0.00385 0.0007 0.6489 0.7926 2.750 0.3043 0.01042 0.00396 0.0010 0.6375 0.8011 3.000 0.3299 0.01041 0.00392 0.0017 0.6127 0.8098 3.250 0.3549 0.01042 0.00387 0.0025 0.5751 0.8182 3.500 0.3802 0.01051 0.00390 0.0032 0.5383 0.8275 3.750 0.4054 0.01063 0.00401 0.0038 0.5027 0.8361 4.000 0.4292 0.01090 0.00411 0.0046 0.4386 0.8459 4.250 0.4472 0.01181 0.00440 0.0060 0.2956 0.8561 4.500 0.4647 0.01292 0.00498 0.0071 0.1721 0.8666 4.750 0.4843 0.01379 0.00550 0.0081 0.0986 0.8778 5.000 0.5061 0.01439 0.00599 0.0089 0.0688 0.8897 5.250 0.5287 0.01489 0.00647 0.0097 0.0554 0.9019 5.500 0.5517 0.01539 0.00700 0.0104 0.0476 0.9148 5.750 0.5760 0.01586 0.00756 0.0108 0.0428 0.9288 6.000 0.6021 0.01642 0.00815 0.0107 0.0383 0.9434 6.250 0.6301 0.01717 0.00896 0.0101 0.0353 0.9585 6.750 0.6930 0.01872 0.01068 0.0072 0.0308 0.9987 7.000 0.7130 0.01948 0.01143 0.0080 0.0292 1.0000 7.250 0.7313 0.02059 0.01252 0.0090 0.0277 1.0000 7.500 0.7531 0.02131 0.01335 0.0096 0.0264 1.0000 7.750 0.7743 0.02225 0.01437 0.0103 0.0254 1.0000 8.000 0.7956 0.02327 0.01547 0.0110 0.0246 1.0000 8.250 0.8170 0.02435 0.01663 0.0116 0.0238 1.0000 8.500 0.8382 0.02547 0.01783 0.0122 0.0230 1.0000 8.750 0.8590 0.02668 0.01915 0.0128 0.0225 1.0000 9.000 0.8792 0.02806 0.02061 0.0134 0.0220 1.0000 9.250 0.8980 0.03009 0.02274 0.0139 0.0213 1.0000 9.500 0.9156 0.03150 0.02443 0.0149 0.0207 1.0000 9.750 0.9313 0.03316 0.02639 0.0159 0.0199 1.0000 10.000 0.9441 0.03535 0.02888 0.0171 0.0195 1.0000 10.250 0.9534 0.03770 0.03155 0.0186 0.0191 1.0000 10.500 0.9590 0.04016 0.03432 0.0202 0.0188 1.0000 10.750 0.9589 0.04278 0.03723 0.0221 0.0185 1.0000 11.000 0.9522 0.04552 0.04022 0.0245 0.0183 1.0000 11.250 0.9429 0.04841 0.04335 0.0263 0.0182 1.0000 11.500 0.9316 0.05158 0.04675 0.0273 0.0180 1.0000 11.750 0.9163 0.05542 0.05081 0.0275 0.0179 1.0000 12.000 0.9013 0.05949 0.05507 0.0268 0.0178 1.0000 12.250 0.8796 0.06492 0.06072 0.0248 0.0177 1.0000 12.500 0.8435 0.07348 0.06954 0.0201 0.0180 1.0000 12.750 0.7936 0.08690 0.08320 0.0110 0.0185 1.0000