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 = 0.200 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -17.750 -0.9636 0.09841 0.09250 -0.0489 1.0000 0.0390 -17.500 -0.9886 0.09026 0.08410 -0.0537 1.0000 0.0392 -17.250 -0.9929 0.08596 0.07974 -0.0560 1.0000 0.0394 -17.000 -0.9977 0.08167 0.07536 -0.0582 1.0000 0.0397 -16.750 -1.0020 0.07755 0.07114 -0.0602 1.0000 0.0400 -16.500 -1.0063 0.07353 0.06702 -0.0621 1.0000 0.0402 -16.250 -1.0101 0.06973 0.06311 -0.0638 1.0000 0.0406 -16.000 -1.0129 0.06616 0.05941 -0.0652 1.0000 0.0409 -15.750 -1.0147 0.06283 0.05596 -0.0665 1.0000 0.0413 -15.500 -1.0153 0.05975 0.05275 -0.0674 1.0000 0.0416 -15.250 -1.0146 0.05688 0.04974 -0.0682 1.0000 0.0420 -15.000 -1.0130 0.05421 0.04691 -0.0687 1.0000 0.0425 -14.750 -1.0102 0.05175 0.04431 -0.0691 1.0000 0.0429 -14.500 -1.0065 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14.250 1.3100 0.06010 0.05311 -0.0139 0.0501 0.8229 14.500 1.3047 0.06324 0.05643 -0.0135 0.0499 0.8268 14.750 1.2969 0.06678 0.06016 -0.0134 0.0498 0.8305 15.000 1.2857 0.07084 0.06442 -0.0137 0.0497 0.8339 15.250 1.2697 0.07564 0.06945 -0.0146 0.0496 0.8372 15.500 1.2482 0.08163 0.07569 -0.0166 0.0495 0.8405 15.750 1.2133 0.09037 0.08477 -0.0208 0.0495 0.8432