XFOIL Version 6.96 Calculated polar for: HQ 0/7 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 = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -9.000 -0.6623 0.09326 0.08662 0.0038 1.0000 0.0358 -8.750 -0.6675 0.08692 0.08034 -0.0013 1.0000 0.0348 -8.500 -0.6795 0.08031 0.07375 -0.0070 1.0000 0.0335 -8.000 -0.6946 0.07056 0.06359 -0.0120 1.0000 0.0316 -7.750 -0.6947 0.06591 0.05876 -0.0129 1.0000 0.0315 -7.500 -0.6922 0.06128 0.05390 -0.0134 1.0000 0.0314 -7.250 -0.6869 0.05672 0.04903 -0.0136 1.0000 0.0313 -7.000 -0.6786 0.05225 0.04416 -0.0135 1.0000 0.0313 -6.750 -0.6670 0.04801 0.03943 -0.0131 1.0000 0.0315 -6.500 -0.6529 0.04377 0.03470 -0.0125 1.0000 0.0320 -6.250 -0.6372 0.03960 0.03018 -0.0120 1.0000 0.0335 -6.000 -0.6178 0.03653 0.02678 -0.0113 1.0000 0.0368 -5.750 -0.5968 0.03384 0.02349 -0.0105 1.0000 0.0459 -5.500 -0.5730 0.03091 0.02027 -0.0096 1.0000 0.0507 -5.250 -0.5464 0.02831 0.01720 -0.0084 1.0000 0.0554 -5.000 -0.5217 0.02617 0.01494 -0.0073 1.0000 0.0615 -4.750 -0.4982 0.02471 0.01335 -0.0065 1.0000 0.0771 -4.500 -0.4749 0.02309 0.01159 -0.0053 1.0000 0.0883 -4.250 -0.4523 0.02153 0.00994 -0.0043 1.0000 0.1078 -4.000 -0.4320 0.01955 0.00840 -0.0038 1.0000 0.1840 -3.750 -0.4202 0.01725 0.00758 -0.0016 1.0000 0.4284 -3.500 -0.4081 0.01651 0.00774 0.0033 1.0000 0.6441 -3.250 -0.3928 0.01646 0.00785 0.0082 1.0000 0.7557 -3.000 -0.3708 0.01656 0.00789 0.0121 1.0000 0.8313 -2.750 -0.3383 0.01649 0.00752 0.0125 1.0000 0.8688 -2.500 -0.3008 0.01635 0.00707 0.0112 1.0000 0.8960 -2.250 -0.2524 0.01629 0.00668 0.0079 1.0000 0.9291 -2.000 -0.1902 0.01615 0.00603 0.0014 1.0000 0.9573 -1.750 -0.1424 0.01585 0.00550 -0.0031 1.0000 0.9747 -1.500 -0.0986 0.01552 0.00501 -0.0071 1.0000 0.9906 -1.250 -0.0671 0.01524 0.00463 -0.0089 1.0000 1.0000 -1.000 -0.0518 0.01502 0.00440 -0.0076 1.0000 1.0000 -0.750 -0.0374 0.01484 0.00419 -0.0060 1.0000 1.0000 -0.500 -0.0240 0.01471 0.00405 -0.0042 1.0000 1.0000 -0.250 -0.0116 0.01463 0.00397 -0.0022 1.0000 1.0000 0.000 0.0000 0.01460 0.00394 0.0000 1.0000 1.0000 0.250 0.0116 0.01463 0.00396 0.0022 1.0000 1.0000 0.500 0.0240 0.01471 0.00405 0.0042 1.0000 1.0000 0.750 0.0375 0.01484 0.00418 0.0060 1.0000 1.0000 1.000 0.0519 0.01502 0.00439 0.0076 1.0000 1.0000 1.250 0.0673 0.01523 0.00463 0.0089 1.0000 1.0000 1.500 0.0987 0.01552 0.00500 0.0071 0.9907 1.0000 1.750 0.1424 0.01584 0.00549 0.0031 0.9747 1.0000 2.000 0.1903 0.01614 0.00602 -0.0014 0.9574 1.0000 2.250 0.2525 0.01629 0.00667 -0.0079 0.9291 1.0000 2.500 0.3009 0.01635 0.00706 -0.0112 0.8961 1.0000 2.750 0.3383 0.01648 0.00752 -0.0125 0.8688 1.0000 3.000 0.3707 0.01655 0.00789 -0.0121 0.8313 1.0000 3.250 0.3927 0.01645 0.00785 -0.0082 0.7558 1.0000 3.500 0.4079 0.01651 0.00774 -0.0033 0.6432 1.0000 3.750 0.4201 0.01725 0.00757 0.0016 0.4287 1.0000 4.000 0.4320 0.01956 0.00840 0.0038 0.1838 1.0000 4.250 0.4524 0.02153 0.00994 0.0043 0.1078 1.0000 4.500 0.4750 0.02309 0.01159 0.0053 0.0881 1.0000 4.750 0.4984 0.02472 0.01336 0.0065 0.0769 1.0000 5.000 0.5219 0.02617 0.01494 0.0073 0.0615 1.0000 5.250 0.5466 0.02832 0.01720 0.0084 0.0554 1.0000 5.500 0.5733 0.03092 0.02028 0.0096 0.0507 1.0000 5.750 0.5970 0.03386 0.02350 0.0105 0.0457 1.0000 6.000 0.6182 0.03653 0.02678 0.0112 0.0366 1.0000 6.250 0.6376 0.03965 0.03023 0.0119 0.0334 1.0000 6.500 0.6535 0.04377 0.03471 0.0124 0.0321 1.0000 6.750 0.6676 0.04806 0.03948 0.0130 0.0315 1.0000 7.000 0.6791 0.05236 0.04428 0.0134 0.0313 1.0000 7.250 0.6876 0.05679 0.04911 0.0134 0.0313 1.0000 7.500 0.6927 0.06142 0.05405 0.0132 0.0313 1.0000 7.750 0.6956 0.06597 0.05883 0.0127 0.0315 1.0000 8.000 0.6959 0.07061 0.06363 0.0118 0.0317 1.0000 8.500 0.6805 0.08045 0.07390 0.0068 0.0335 1.0000 8.750 0.6685 0.08706 0.08048 0.0011 0.0347 1.0000 9.000 0.6634 0.09340 0.08676 -0.0041 0.0358 1.0000