XFOIL Version 6.96 Calculated polar for: FX 38-153 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 ------ -------- --------- --------- -------- -------- -------- -12.250 -0.5175 0.04621 0.04269 -0.1277 0.9617 0.0053 -12.000 -0.5410 0.04111 0.03723 -0.1268 0.9585 0.0053 -11.750 -0.5523 0.03749 0.03330 -0.1254 0.9560 0.0053 -11.500 -0.5569 0.03471 0.03025 -0.1239 0.9536 0.0053 -11.250 -0.5586 0.03253 0.02784 -0.1218 0.9505 0.0053 -11.000 -0.5557 0.03055 0.02563 -0.1199 0.9474 0.0054 -10.750 -0.5483 0.02882 0.02368 -0.1184 0.9449 0.0054 -10.500 -0.5373 0.02724 0.02191 -0.1171 0.9430 0.0055 -10.250 -0.5233 0.02582 0.02028 -0.1161 0.9414 0.0056 -10.000 -0.5068 0.02455 0.01884 -0.1152 0.9399 0.0057 -9.750 -0.4915 0.02351 0.01766 -0.1139 0.9378 0.0058 -9.500 -0.4756 0.02258 0.01660 -0.1126 0.9356 0.0059 -9.250 -0.4583 0.02173 0.01563 -0.1114 0.9334 0.0060 -9.000 -0.4396 0.02090 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-0.0583 0.0130 0.8212 10.500 1.2565 0.02657 0.01889 -0.0572 0.0129 0.8222 10.750 1.2719 0.02736 0.01976 -0.0559 0.0126 0.8232 11.000 1.2862 0.02821 0.02069 -0.0546 0.0122 0.8242 11.250 1.2998 0.02914 0.02171 -0.0533 0.0117 0.8252 11.500 1.3121 0.03017 0.02282 -0.0518 0.0113 0.8263 11.750 1.3234 0.03129 0.02402 -0.0503 0.0109 0.8273 12.000 1.3339 0.03248 0.02529 -0.0488 0.0106 0.8284 12.250 1.3433 0.03377 0.02668 -0.0472 0.0103 0.8296 12.500 1.3507 0.03526 0.02825 -0.0456 0.0101 0.8308 12.750 1.3628 0.03639 0.02947 -0.0445 0.0099 0.8319 13.000 1.3737 0.03758 0.03076 -0.0433 0.0097 0.8327 13.250 1.3835 0.03888 0.03216 -0.0420 0.0095 0.8336 13.500 1.3928 0.04026 0.03364 -0.0409 0.0092 0.8345 13.750 1.4012 0.04177 0.03526 -0.0397 0.0090 0.8354 14.000 1.4088 0.04339 0.03699 -0.0386 0.0087 0.8363 14.250 1.4157 0.04511 0.03881 -0.0376 0.0083 0.8372 14.500 1.4218 0.04696 0.04077 -0.0367 0.0079 0.8382 14.750 1.4266 0.04901 0.04291 -0.0358 0.0076 0.8393 15.000 1.4318 0.05109 0.04509 -0.0351 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