XFOIL Version 6.96 Calculated polar for: NACA 66-209 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -9.500 -0.5195 0.08378 0.08217 -0.0360 1.0000 0.0090 -9.250 -0.5243 0.07875 0.07718 -0.0394 1.0000 0.0091 -9.000 -0.5378 0.07205 0.07050 -0.0458 1.0000 0.0087 -8.750 -0.5532 0.06818 0.06662 -0.0473 1.0000 0.0092 -8.500 -0.5709 0.06404 0.06244 -0.0470 0.9988 0.0088 -8.250 -0.5606 0.05848 0.05676 -0.0519 0.9882 0.0095 -8.000 -0.5456 0.05223 0.05033 -0.0568 0.9830 0.0099 -7.750 -0.5193 0.04715 0.04500 -0.0590 0.9760 0.0113 -6.750 -0.4838 0.01378 0.01050 -0.0506 0.9172 0.0101 -6.500 -0.4759 0.00892 0.00506 -0.0480 0.9114 0.0096 -6.250 -0.4575 0.00679 0.00256 -0.0468 0.9063 0.0103 -6.000 -0.4364 0.00534 0.00077 -0.0457 0.9019 0.0110 -5.750 -0.4132 0.01582 0.01087 -0.0488 0.9083 0.0115 -5.500 -0.3929 0.01268 0.00740 -0.0475 0.9034 0.0122 -5.250 -0.3682 0.01200 0.00665 -0.0471 0.8991 0.0129 -5.000 -0.3425 0.01156 0.00617 -0.0469 0.8945 0.0137 -4.750 -0.3163 0.01133 0.00590 -0.0468 0.8903 0.0151 -4.500 -0.2910 0.01078 0.00527 -0.0464 0.8864 0.0158 -4.250 -0.2651 0.01036 0.00481 -0.0461 0.8823 0.0165 -4.000 -0.2388 0.01008 0.00448 -0.0460 0.8781 0.0170 -3.750 -0.2168 0.00901 0.00328 -0.0450 0.8740 0.0184 -3.500 -0.1908 0.00867 0.00293 -0.0448 0.8702 0.0206 -3.250 -0.1641 0.00842 0.00266 -0.0448 0.8664 0.0227 -3.000 -0.1373 0.00822 0.00242 -0.0447 0.8629 0.0245 -2.750 -0.1106 0.00796 0.00209 -0.0446 0.8596 0.0269 -2.500 -0.0839 0.00770 0.00181 -0.0445 0.8564 0.0333 -2.250 -0.0565 0.00755 0.00165 -0.0445 0.8529 0.0384 -2.000 -0.0301 0.00721 0.00146 -0.0445 0.8494 0.0852 -1.750 -0.0092 0.00594 0.00119 -0.0441 0.8460 0.3967 -1.500 0.0114 0.00483 0.00106 -0.0433 0.8427 0.6918 -1.250 0.0382 0.00472 0.00110 -0.0431 0.8396 0.7436 -1.000 0.0655 0.00470 0.00114 -0.0430 0.8366 0.7729 -0.750 0.0930 0.00472 0.00117 -0.0429 0.8337 0.7916 -0.500 0.1206 0.00476 0.00122 -0.0428 0.8307 0.8057 -0.250 0.1483 0.00477 0.00126 -0.0428 0.8271 0.8148 0.000 0.1759 0.00477 0.00128 -0.0428 0.8223 0.8231 0.250 0.2033 0.00483 0.00131 -0.0426 0.8176 0.8331 0.500 0.2309 0.00481 0.00133 -0.0426 0.8122 0.8373 0.750 0.2583 0.00481 0.00128 -0.0425 0.8021 0.8406 1.000 0.2855 0.00479 0.00124 -0.0423 0.7891 0.8439 1.250 0.3133 0.00478 0.00124 -0.0424 0.7786 0.8470 1.500 0.3404 0.00478 0.00123 -0.0422 0.7643 0.8499 1.750 0.3672 0.00481 0.00123 -0.0420 0.7453 0.8531 2.000 0.3934 0.00489 0.00124 -0.0417 0.7132 0.8568 2.250 0.4143 0.00536 0.00129 -0.0403 0.5962 0.8608 2.500 0.4267 0.00670 0.00170 -0.0378 0.3602 0.8650 2.750 0.4426 0.00790 0.00211 -0.0362 0.1538 0.8696 3.000 0.4636 0.00869 0.00244 -0.0352 0.0417 0.8740 3.250 0.4886 0.00896 0.00270 -0.0348 0.0319 0.8778 3.500 0.5145 0.00914 0.00293 -0.0345 0.0282 0.8820 3.750 0.5399 0.00942 0.00322 -0.0342 0.0235 0.8866 4.000 0.5633 0.00990 0.00378 -0.0333 0.0202 0.8911 4.250 0.5882 0.01014 0.00407 -0.0329 0.0191 0.8956 4.500 0.6130 0.01047 0.00443 -0.0324 0.0179 0.9006 4.750 0.6370 0.01082 0.00484 -0.0318 0.0167 0.9053 5.000 0.6605 0.01118 0.00523 -0.0311 0.0153 0.9104 5.250 0.6799 0.01215 0.00629 -0.0296 0.0141 0.9166 5.500 0.6976 0.01341 0.00768 -0.0278 0.0136 0.9228 5.750 0.7214 0.01388 0.00821 -0.0271 0.0133 0.9290 6.000 0.7438 0.01445 0.00887 -0.0261 0.0130 0.9351 6.250 0.7665 0.01520 0.00970 -0.0253 0.0126 0.9421 6.500 0.7885 0.01569 0.01029 -0.0242 0.0117 0.9497 6.750 0.8105 0.01612 0.01079 -0.0232 0.0110 0.9590 7.500 0.8900 0.01915 0.01415 -0.0239 0.0095 1.0000 8.000 0.9255 0.02568 0.02145 -0.0212 0.0086 1.0000 8.250 0.9449 0.02764 0.02366 -0.0200 0.0080 1.0000 8.500 0.9533 0.03237 0.02884 -0.0172 0.0074 1.0000 8.750 0.9417 0.04101 0.03811 -0.0119 0.0076 1.0000 9.000 0.9332 0.04710 0.04456 -0.0080 0.0078 1.0000 9.250 0.9229 0.05230 0.05002 -0.0047 0.0081 1.0000