XFOIL Version 6.96 Calculated polar for: NASA/LANGLEY RC12-64C AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.250 -0.4526 0.10071 0.09360 -0.0098 1.0000 0.3504 -8.000 -0.4530 0.09799 0.09096 -0.0086 1.0000 0.3675 -7.500 -0.4436 0.09212 0.08522 -0.0054 1.0000 0.4033 -5.750 -0.6265 0.05287 0.04459 -0.0118 1.0000 0.1311 -5.500 -0.6208 0.04935 0.04045 -0.0087 1.0000 0.1194 -5.250 -0.6111 0.04599 0.03667 -0.0061 1.0000 0.1124 -5.000 -0.5979 0.04342 0.03351 -0.0033 1.0000 0.1063 -4.750 -0.5818 0.04080 0.03064 -0.0014 1.0000 0.1031 -4.500 -0.5637 0.03833 0.02774 0.0006 1.0000 0.0993 -4.250 -0.5436 0.03657 0.02529 0.0029 1.0000 0.0957 -4.000 -0.5211 0.03466 0.02317 0.0041 1.0000 0.0945 -3.750 -0.4973 0.03294 0.02128 0.0051 1.0000 0.0937 -3.500 -0.4721 0.03146 0.01964 0.0061 1.0000 0.0934 -3.250 -0.4458 0.03027 0.01830 0.0069 1.0000 0.0943 -3.000 -0.4186 0.02918 0.01716 0.0076 1.0000 0.0963 -2.750 -0.3928 0.02823 0.01626 0.0082 1.0000 0.0996 -2.500 -0.3695 0.02749 0.01546 0.0093 1.0000 0.1033 -2.250 -0.3484 0.02686 0.01471 0.0107 1.0000 0.1073 -2.000 -0.0694 0.02452 0.01556 -0.0251 1.0000 1.0000 -1.750 -0.0669 0.02444 0.01533 -0.0217 1.0000 1.0000 -1.500 -0.0646 0.02439 0.01516 -0.0182 1.0000 1.0000 -1.250 -0.0625 0.02437 0.01501 -0.0147 1.0000 1.0000 -1.000 -0.0603 0.02439 0.01492 -0.0113 1.0000 1.0000 -0.750 -0.0578 0.02443 0.01485 -0.0078 1.0000 1.0000 -0.500 -0.0550 0.02451 0.01483 -0.0045 1.0000 1.0000 -0.250 -0.0516 0.02462 0.01484 -0.0013 1.0000 1.0000 0.000 -0.0475 0.02476 0.01489 0.0018 1.0000 1.0000 0.250 -0.0426 0.02494 0.01497 0.0047 1.0000 1.0000 0.500 -0.0368 0.02517 0.01510 0.0075 1.0000 1.0000 0.750 -0.0300 0.02543 0.01528 0.0100 1.0000 1.0000 1.000 -0.0223 0.02574 0.01551 0.0123 1.0000 1.0000 1.250 -0.0138 0.02610 0.01579 0.0144 1.0000 1.0000 1.500 -0.0046 0.02651 0.01612 0.0164 1.0000 1.0000 1.750 0.0055 0.02697 0.01652 0.0181 1.0000 1.0000 2.000 0.0164 0.02749 0.01699 0.0196 1.0000 1.0000 2.250 0.0280 0.02807 0.01752 0.0209 1.0000 1.0000 2.500 0.0401 0.02872 0.01813 0.0219 1.0000 1.0000 2.750 0.0813 0.02996 0.01937 0.0174 0.9894 1.0000 3.000 0.1372 0.03143 0.02086 0.0102 0.9707 1.0000 3.250 0.1885 0.03266 0.02214 0.0042 0.9500 1.0000 3.500 0.2457 0.03385 0.02342 -0.0024 0.9286 1.0000 3.750 0.2878 0.03459 0.02427 -0.0059 0.9043 1.0000 4.000 0.3557 0.03517 0.02504 -0.0131 0.8787 1.0000 4.250 0.4019 0.03530 0.02535 -0.0160 0.8490 1.0000 4.500 0.4657 0.03482 0.02514 -0.0207 0.8178 1.0000 4.750 0.6056 0.03032 0.02129 -0.0323 0.7721 1.0000 5.000 0.6684 0.02711 0.01842 -0.0322 0.7258 1.0000 5.250 0.7103 0.02452 0.01594 -0.0289 0.6649 1.0000 5.500 0.7315 0.02324 0.01437 -0.0232 0.5624 1.0000 5.750 0.7343 0.02416 0.01403 -0.0161 0.4085 1.0000 6.000 0.7426 0.02622 0.01513 -0.0121 0.3181 1.0000 6.250 0.7655 0.02808 0.01647 -0.0110 0.2705 1.0000 6.500 0.7915 0.02970 0.01795 -0.0103 0.2417 1.0000 6.750 0.8201 0.03146 0.01946 -0.0102 0.2215 1.0000 7.000 0.8434 0.03308 0.02116 -0.0092 0.2065 1.0000 7.250 0.8644 0.03482 0.02307 -0.0078 0.1942 1.0000 7.500 0.8856 0.03683 0.02516 -0.0067 0.1844 1.0000 7.750 0.9048 0.03872 0.02718 -0.0052 0.1753 1.0000 8.000 0.9206 0.04112 0.02984 -0.0033 0.1689 1.0000 8.250 0.9325 0.04334 0.03240 -0.0007 0.1627 1.0000 8.500 0.9525 0.04599 0.03500 0.0002 0.1570 1.0000 8.750 0.9556 0.04879 0.03829 0.0037 0.1541 1.0000 9.000 0.9554 0.05169 0.04166 0.0073 0.1514 1.0000 9.250 0.9541 0.05490 0.04526 0.0107 0.1497 1.0000 9.500 0.9492 0.05847 0.04920 0.0141 0.1493 1.0000 9.750 0.9385 0.06229 0.05334 0.0176 0.1494 1.0000 10.000 0.9218 0.06638 0.05773 0.0210 0.1501 1.0000 10.250 0.9004 0.07073 0.06230 0.0240 0.1514 1.0000 10.500 0.8766 0.07517 0.06688 0.0267 0.1530 1.0000 10.750 0.8557 0.07990 0.07169 0.0284 0.1545 1.0000 11.000 0.8437 0.08507 0.07690 0.0289 0.1558 1.0000