XFOIL Version 6.94 Calculated polar for: GRUMMAN K-2 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.0534 0.01039 0.00648 -0.0521 0.9715 0.7746 0.500 0.1322 0.01009 0.00622 -0.0558 0.9579 0.7824 1.000 0.2002 0.00998 0.00618 -0.0571 0.9457 0.7915 2.000 0.3328 0.00912 0.00545 -0.0590 0.8918 0.8018 2.500 0.3973 0.01013 0.00518 -0.0593 0.5822 0.8042 3.000 0.4392 0.01168 0.00555 -0.0571 0.3222 0.8110 3.500 0.4836 0.01266 0.00597 -0.0547 0.1919 0.8139 4.000 0.5357 0.01332 0.00636 -0.0541 0.1375 0.8183 4.500 0.5906 0.01381 0.00674 -0.0540 0.1143 0.8227 5.000 0.6417 0.01434 0.00721 -0.0530 0.0951 0.8260 5.500 0.7006 0.01483 0.00761 -0.0539 0.0710 0.8313 6.000 0.7494 0.01547 0.00819 -0.0523 0.0536 0.8339 6.500 0.8011 0.01631 0.00900 -0.0516 0.0445 0.8379 7.000 0.8541 0.01723 0.00991 -0.0513 0.0379 0.8420 7.500 0.9006 0.01833 0.01110 -0.0494 0.0344 0.8451 8.000 0.9512 0.01973 0.01254 -0.0487 0.0315 0.8492 8.500 0.9987 0.02108 0.01395 -0.0474 0.0293 0.8524 9.000 1.0442 0.02271 0.01575 -0.0457 0.0280 0.8555 9.500 1.0940 0.02434 0.01753 -0.0451 0.0255 0.8598 10.000 1.1343 0.02671 0.02016 -0.0427 0.0242 0.8624 10.500 1.1746 0.02871 0.02249 -0.0404 0.0233 0.8660 11.000 1.2129 0.03179 0.02593 -0.0384 0.0222 0.8698 11.500 1.2396 0.03478 0.02915 -0.0348 0.0210 0.8725 12.000 1.2523 0.03927 0.03426 -0.0296 0.0208 0.8763 12.500 1.2433 0.04435 0.03997 -0.0224 0.0205 0.8801 13.000 1.2213 0.04955 0.04571 -0.0155 0.0202 0.8838 13.500 1.1931 0.05696 0.05349 -0.0125 0.0207 0.8874