XFOIL Version 6.96 Calculated polar for: S1010 HPV 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 = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -9.500 -0.5767 0.09424 0.09095 0.0091 1.0000 0.0594 -9.250 -0.5864 0.08850 0.08524 0.0055 1.0000 0.0611 -9.000 -0.6729 0.09063 0.08721 0.0073 1.0000 0.0563 -8.750 -0.6770 0.08591 0.08250 0.0042 1.0000 0.0575 -8.500 -0.6859 0.08067 0.07734 -0.0001 1.0000 0.0585 -8.250 -0.7587 0.05187 0.04793 -0.0159 1.0000 0.0278 -8.000 -0.7635 0.04541 0.04111 -0.0160 1.0000 0.0273 -7.500 -0.7543 0.03503 0.02955 -0.0141 1.0000 0.0287 -7.250 -0.7448 0.02984 0.02381 -0.0129 1.0000 0.0297 -7.000 -0.7275 0.02671 0.02029 -0.0117 1.0000 0.0308 -6.750 -0.7070 0.02452 0.01780 -0.0107 1.0000 0.0327 -6.500 -0.6845 0.02293 0.01590 -0.0097 1.0000 0.0360 -6.250 -0.6611 0.02154 0.01413 -0.0086 1.0000 0.0392 -6.000 -0.6393 0.01914 0.01156 -0.0077 1.0000 0.0429 -5.750 -0.6150 0.01832 0.01059 -0.0069 1.0000 0.0485 -5.500 -0.5918 0.01679 0.00896 -0.0059 1.0000 0.0544 -5.250 -0.5673 0.01601 0.00804 -0.0050 1.0000 0.0626 -5.000 -0.5437 0.01519 0.00722 -0.0042 1.0000 0.0753 -4.750 -0.5199 0.01450 0.00650 -0.0034 1.0000 0.0888 -4.500 -0.4961 0.01391 0.00589 -0.0026 1.0000 0.1026 -4.250 -0.4726 0.01334 0.00537 -0.0018 1.0000 0.1178 -4.000 -0.4492 0.01282 0.00491 -0.0009 1.0000 0.1336 -3.750 -0.4258 0.01232 0.00446 0.0000 1.0000 0.1505 -3.500 -0.4023 0.01188 0.00405 0.0009 1.0000 0.1713 -3.250 -0.3796 0.01134 0.00363 0.0018 1.0000 0.1945 -3.000 -0.3568 0.01085 0.00329 0.0028 1.0000 0.2256 -2.750 -0.3348 0.01028 0.00299 0.0037 1.0000 0.2735 -2.500 -0.3142 0.00956 0.00273 0.0049 1.0000 0.3658 -2.250 -0.2981 0.00849 0.00253 0.0070 1.0000 0.5528 -2.000 -0.2868 0.00769 0.00271 0.0116 1.0000 0.7963 -1.750 -0.2627 0.00767 0.00286 0.0137 1.0000 0.8938 -1.500 -0.2131 0.00786 0.00301 0.0103 1.0000 0.9493 -1.250 -0.1492 0.00802 0.00303 0.0034 1.0000 0.9764 -1.000 -0.0836 0.00804 0.00294 -0.0041 1.0000 0.9925 -0.750 -0.0378 0.00795 0.00280 -0.0080 1.0000 1.0000 -0.500 -0.0239 0.00788 0.00271 -0.0056 1.0000 1.0000 -0.250 -0.0114 0.00784 0.00267 -0.0029 1.0000 1.0000 0.000 0.0000 0.00783 0.00266 0.0000 1.0000 1.0000 0.250 0.0114 0.00784 0.00267 0.0029 1.0000 1.0000 0.500 0.0239 0.00788 0.00271 0.0056 1.0000 1.0000 0.750 0.0379 0.00795 0.00280 0.0080 1.0000 1.0000 1.000 0.0840 0.00804 0.00294 0.0041 0.9924 1.0000 1.250 0.1486 0.00802 0.00303 -0.0033 0.9767 1.0000 1.500 0.2133 0.00786 0.00301 -0.0103 0.9491 1.0000 1.750 0.2627 0.00767 0.00286 -0.0137 0.8942 1.0000 2.000 0.2869 0.00769 0.00271 -0.0116 0.7965 1.0000 2.250 0.2982 0.00849 0.00253 -0.0070 0.5527 1.0000 2.500 0.3143 0.00956 0.00273 -0.0049 0.3661 1.0000 2.750 0.3349 0.01028 0.00299 -0.0038 0.2738 1.0000 3.000 0.3569 0.01085 0.00329 -0.0028 0.2257 1.0000 3.250 0.3797 0.01134 0.00363 -0.0018 0.1946 1.0000 3.500 0.4023 0.01188 0.00405 -0.0009 0.1713 1.0000 3.750 0.4258 0.01232 0.00446 0.0000 0.1505 1.0000 4.000 0.4492 0.01282 0.00491 0.0009 0.1336 1.0000 4.250 0.4727 0.01334 0.00537 0.0018 0.1178 1.0000 4.500 0.4961 0.01392 0.00589 0.0026 0.1025 1.0000 4.750 0.5199 0.01451 0.00650 0.0034 0.0888 1.0000 5.000 0.5437 0.01519 0.00722 0.0042 0.0753 1.0000 5.250 0.5674 0.01600 0.00804 0.0050 0.0625 1.0000 5.500 0.5919 0.01679 0.00896 0.0059 0.0544 1.0000 5.750 0.6150 0.01832 0.01058 0.0069 0.0486 1.0000 6.000 0.6394 0.01914 0.01155 0.0077 0.0429 1.0000 6.250 0.6611 0.02153 0.01412 0.0086 0.0392 1.0000 6.500 0.6845 0.02296 0.01594 0.0097 0.0361 1.0000 6.750 0.7070 0.02453 0.01781 0.0107 0.0327 1.0000 7.000 0.7275 0.02673 0.02032 0.0117 0.0308 1.0000 7.250 0.7449 0.02984 0.02380 0.0129 0.0297 1.0000 7.500 0.7555 0.03469 0.02918 0.0141 0.0289 1.0000 7.750 0.7595 0.04031 0.03546 0.0153 0.0279 1.0000 8.000 0.7636 0.04538 0.04104 0.0160 0.0276 1.0000 8.250 0.7591 0.05181 0.04784 0.0159 0.0279 1.0000 8.500 0.6863 0.08070 0.07738 0.0000 0.0584 1.0000 8.750 0.6775 0.08597 0.08259 -0.0043 0.0577 1.0000