XFOIL Version 6.96 Calculated polar for: NACA 0006 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 ------ -------- --------- --------- -------- -------- -------- -9.250 -0.7553 0.08620 0.08416 0.0189 1.0000 0.0038 -9.000 -0.7635 0.07934 0.07734 0.0133 1.0000 0.0038 -8.750 -0.7723 0.07106 0.06905 0.0035 1.0000 0.0037 -8.500 -0.7784 0.06178 0.05964 -0.0037 1.0000 0.0037 -8.250 -0.7851 0.05186 0.04946 -0.0078 1.0000 0.0037 -8.000 -0.8037 0.03568 0.03252 -0.0089 1.0000 0.0036 -7.750 -0.7984 0.02789 0.02390 -0.0078 1.0000 0.0038 -7.500 -0.7839 0.02371 0.01914 -0.0068 1.0000 0.0041 -7.250 -0.7629 0.02183 0.01701 -0.0063 1.0000 0.0043 -7.000 -0.7412 0.01988 0.01474 -0.0056 1.0000 0.0045 -6.750 -0.7184 0.01823 0.01283 -0.0050 1.0000 0.0047 -6.500 -0.6948 0.01680 0.01115 -0.0043 1.0000 0.0050 -6.250 -0.6708 0.01555 0.00965 -0.0037 1.0000 0.0054 -6.000 -0.6464 0.01447 0.00839 -0.0031 1.0000 0.0058 -5.750 -0.6215 0.01362 0.00739 -0.0026 1.0000 0.0062 -5.500 -0.5972 0.01261 0.00626 -0.0021 1.0000 0.0072 -5.250 -0.5715 0.01215 0.00575 -0.0018 1.0000 0.0083 -5.000 -0.5460 0.01154 0.00505 -0.0013 1.0000 0.0094 -4.750 -0.5202 0.01105 0.00445 -0.0009 1.0000 0.0103 -4.500 -0.4948 0.01039 0.00373 -0.0004 1.0000 0.0150 -4.250 -0.4687 0.01007 0.00339 -0.0001 1.0000 0.0197 -4.000 -0.4425 0.00977 0.00310 0.0002 1.0000 0.0277 -3.750 -0.4162 0.00955 0.00282 0.0005 1.0000 0.0323 -3.500 -0.3902 0.00928 0.00254 0.0008 1.0000 0.0373 -3.250 -0.3641 0.00906 0.00229 0.0011 1.0000 0.0420 -3.000 -0.3382 0.00881 0.00205 0.0015 1.0000 0.0477 -2.750 -0.3124 0.00855 0.00182 0.0019 1.0000 0.0605 -2.250 -0.2620 0.00777 0.00145 0.0026 1.0000 0.1646 -2.000 -0.2368 0.00741 0.00131 0.0030 1.0000 0.2270 -1.750 -0.2116 0.00703 0.00118 0.0033 1.0000 0.2997 -1.250 -0.1501 0.00622 0.00102 0.0014 0.9927 0.4780 -1.000 -0.1164 0.00586 0.00096 -0.0001 0.9846 0.5625 -0.750 -0.0847 0.00541 0.00094 -0.0010 0.9737 0.6727 -0.500 -0.0539 0.00509 0.00094 -0.0016 0.9591 0.7544 -0.250 -0.0256 0.00481 0.00097 -0.0012 0.9394 0.8430 0.000 0.0000 0.00470 0.00101 0.0000 0.9073 0.9077 0.250 0.0257 0.00480 0.00097 0.0012 0.8434 0.9394 0.500 0.0539 0.00509 0.00094 0.0016 0.7544 0.9591 0.750 0.0847 0.00541 0.00094 0.0010 0.6726 0.9738 1.000 0.1164 0.00586 0.00096 0.0001 0.5627 0.9846 1.250 0.1500 0.00622 0.00102 -0.0014 0.4786 0.9928 1.500 0.1841 0.00667 0.00110 -0.0031 0.3761 0.9987 1.750 0.2115 0.00703 0.00118 -0.0033 0.2996 1.0000 2.000 0.2367 0.00741 0.00131 -0.0029 0.2272 1.0000 2.250 0.2618 0.00777 0.00145 -0.0026 0.1651 1.0000 2.750 0.3122 0.00855 0.00182 -0.0018 0.0607 1.0000 3.000 0.3380 0.00880 0.00205 -0.0015 0.0478 1.0000 3.250 0.3640 0.00906 0.00229 -0.0011 0.0420 1.0000 3.500 0.3901 0.00928 0.00254 -0.0008 0.0373 1.0000 3.750 0.4161 0.00955 0.00282 -0.0004 0.0323 1.0000 4.000 0.4424 0.00977 0.00309 -0.0002 0.0277 1.0000 4.250 0.4686 0.01007 0.00339 0.0001 0.0196 1.0000 4.500 0.4947 0.01039 0.00373 0.0005 0.0150 1.0000 4.750 0.5202 0.01105 0.00445 0.0009 0.0103 1.0000 5.000 0.5460 0.01154 0.00505 0.0013 0.0094 1.0000 5.250 0.5715 0.01214 0.00575 0.0018 0.0083 1.0000 5.500 0.5973 0.01260 0.00624 0.0021 0.0072 1.0000 5.750 0.6216 0.01361 0.00738 0.0026 0.0062 1.0000 6.000 0.6465 0.01447 0.00839 0.0031 0.0058 1.0000 6.250 0.6709 0.01554 0.00964 0.0037 0.0054 1.0000 6.500 0.6949 0.01681 0.01117 0.0043 0.0050 1.0000 6.750 0.7185 0.01824 0.01284 0.0049 0.0047 1.0000 7.000 0.7414 0.01988 0.01474 0.0056 0.0045 1.0000 7.250 0.7631 0.02184 0.01701 0.0062 0.0043 1.0000 7.500 0.7841 0.02372 0.01916 0.0068 0.0041 1.0000 7.750 0.7986 0.02793 0.02395 0.0077 0.0038 1.0000 8.000 0.8037 0.03582 0.03267 0.0089 0.0036 1.0000 8.250 0.7853 0.05197 0.04957 0.0077 0.0037 1.0000 8.500 0.7787 0.06187 0.05972 0.0036 0.0037 1.0000 8.750 0.7726 0.07120 0.06919 -0.0038 0.0037 1.0000 9.000 0.7641 0.07942 0.07742 -0.0135 0.0038 1.0000 9.250 0.7557 0.08640 0.08435 -0.0192 0.0038 1.0000