XFOIL Version 6.96 Calculated polar for: NPL 9510 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 -1.0373 0.15102 0.14781 0.0038 1.0000 0.0158 -19.500 -1.0530 0.14321 0.13988 -0.0007 1.0000 0.0159 -19.250 -1.0676 0.13586 0.13243 -0.0048 1.0000 0.0160 -19.000 -1.0814 0.12887 0.12533 -0.0086 1.0000 0.0161 -18.750 -1.0948 0.12212 0.11847 -0.0123 1.0000 0.0162 -18.500 -1.1072 0.11566 0.11190 -0.0157 1.0000 0.0163 -18.250 -1.1188 0.10953 0.10566 -0.0189 1.0000 0.0164 -18.000 -1.1294 0.10368 0.09970 -0.0219 1.0000 0.0165 -17.750 -1.1392 0.09807 0.09397 -0.0247 1.0000 0.0166 -17.500 -1.1481 0.09273 0.08852 -0.0274 1.0000 0.0168 -17.250 -1.1561 0.08766 0.08334 -0.0299 1.0000 0.0169 -17.000 -1.1629 0.08286 0.07842 -0.0322 1.0000 0.0170 -16.750 -1.1688 0.07833 0.07379 -0.0343 1.0000 0.0171 -16.500 -1.1735 0.07406 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0.0355 0.5631 2.500 0.4982 0.01256 0.00427 -0.0822 0.0339 0.5665 2.750 0.5254 0.01287 0.00457 -0.0820 0.0326 0.5699 3.000 0.5529 0.01314 0.00485 -0.0817 0.0317 0.5735 3.250 0.5803 0.01339 0.00514 -0.0815 0.0306 0.5769 3.500 0.6077 0.01365 0.00542 -0.0813 0.0295 0.5803 3.750 0.6348 0.01393 0.00571 -0.0811 0.0285 0.5837 4.000 0.6612 0.01436 0.00613 -0.0807 0.0276 0.5873 4.250 0.6882 0.01470 0.00651 -0.0804 0.0270 0.5909 4.500 0.7149 0.01506 0.00693 -0.0800 0.0263 0.5944 4.750 0.7416 0.01541 0.00733 -0.0797 0.0254 0.5981 5.000 0.7683 0.01575 0.00769 -0.0794 0.0247 0.6019 5.250 0.7948 0.01611 0.00806 -0.0790 0.0240 0.6056 5.500 0.8206 0.01665 0.00862 -0.0786 0.0234 0.6093 5.750 0.8469 0.01710 0.00917 -0.0782 0.0228 0.6131 6.000 0.8729 0.01760 0.00976 -0.0777 0.0222 0.6172 6.250 0.8987 0.01811 0.01035 -0.0772 0.0217 0.6214 6.500 0.9246 0.01861 0.01090 -0.0768 0.0212 0.6254 6.750 0.9502 0.01909 0.01145 -0.0763 0.0207 0.6293 7.000 0.9757 0.01959 0.01200 -0.0758 0.0204 0.6338 7.250 1.0005 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