XFOIL Version 6.96 Calculated polar for: FX 73-CL1-152 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 -0.0429 0.09916 0.09471 -0.0701 0.5414 0.0412 -9.750 -0.0337 0.09690 0.09240 -0.0709 0.5338 0.0421 -9.500 -0.0578 0.09030 0.08585 -0.0770 0.5336 0.0448 -9.250 -0.0458 0.08805 0.08358 -0.0770 0.5264 0.0450 -9.000 -0.0308 0.08649 0.08198 -0.0766 0.5191 0.0453 -8.750 -0.0180 0.08472 0.08014 -0.0766 0.5124 0.0456 -8.500 -0.2796 0.03080 0.02570 -0.1175 0.5468 0.0397 -8.250 -0.2730 0.02655 0.02093 -0.1189 0.5394 0.0400 -8.000 -0.2539 0.02486 0.01903 -0.1188 0.5311 0.0404 -7.750 -0.2309 0.02385 0.01793 -0.1187 0.5236 0.0409 -7.500 -0.2071 0.02312 0.01708 -0.1184 0.5161 0.0414 -7.250 -0.1832 0.02224 0.01605 -0.1182 0.5098 0.0419 -7.000 -0.1592 0.02117 0.01478 -0.1180 0.5038 0.0426 -6.750 -0.1349 0.02017 0.01354 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13.000 1.8704 0.02062 0.01510 -0.0793 0.2936 1.0000 13.250 1.8823 0.02124 0.01568 -0.0774 0.2897 1.0000 13.500 1.8918 0.02190 0.01646 -0.0754 0.2869 1.0000 13.750 1.9012 0.02264 0.01730 -0.0736 0.2834 1.0000 14.000 1.9102 0.02351 0.01823 -0.0720 0.2800 1.0000 14.250 1.9176 0.02458 0.01929 -0.0705 0.2765 1.0000 14.500 1.9246 0.02580 0.02060 -0.0692 0.2729 1.0000 14.750 1.9307 0.02721 0.02212 -0.0682 0.2690 1.0000 15.000 1.9348 0.02887 0.02385 -0.0674 0.2647 1.0000 15.250 1.9352 0.03095 0.02592 -0.0665 0.2600 1.0000 15.500 1.9384 0.03305 0.02817 -0.0662 0.2561 1.0000 15.750 1.9379 0.03553 0.03073 -0.0660 0.2516 1.0000 16.000 1.9345 0.03830 0.03353 -0.0656 0.2478 1.0000 16.250 1.9307 0.04119 0.03651 -0.0654 0.2442 1.0000 16.500 1.9260 0.04430 0.03975 -0.0654 0.2403 1.0000 16.750 1.9168 0.04784 0.04336 -0.0653 0.2361 1.0000 17.000 1.9040 0.05173 0.04726 -0.0653 0.2320 1.0000 17.250 1.8936 0.05569 0.05137 -0.0656 0.2278 1.0000 17.500 1.8776 0.06029 0.05605 -0.0659 0.2227 1.0000 17.750 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