XFOIL Version 6.96 Calculated polar for: GOE 670 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -13.750 -1.0542 0.02898 0.02604 -0.0872 0.9931 0.0099 -13.500 -1.0409 0.02601 0.02283 -0.0878 0.9899 0.0105 -13.250 -1.0210 0.02403 0.02065 -0.0885 0.9877 0.0110 -13.000 -0.9984 0.02235 0.01880 -0.0892 0.9861 0.0119 -12.750 -0.9754 0.02103 0.01736 -0.0897 0.9840 0.0129 -12.500 -0.9524 0.02003 0.01626 -0.0896 0.9809 0.0138 -12.250 -0.9261 0.01913 0.01524 -0.0902 0.9786 0.0148 -12.000 -0.8974 0.01843 0.01450 -0.0910 0.9770 0.0158 -11.750 -0.8669 0.01793 0.01396 -0.0921 0.9758 0.0169 -11.500 -0.8405 0.01744 0.01342 -0.0922 0.9728 0.0179 -11.250 -0.8131 0.01699 0.01288 -0.0925 0.9699 0.0187 -11.000 -0.7844 0.01646 0.01225 -0.0931 0.9677 0.0191 -10.750 -0.7548 0.01588 0.01158 -0.0940 0.9658 0.0197 -10.500 -0.7278 0.01546 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0.2069 1.0000 4.250 0.8410 0.00881 0.00300 -0.0801 0.2027 1.0000 4.500 0.8628 0.00896 0.00313 -0.0789 0.1969 1.0000 4.750 0.8847 0.00913 0.00327 -0.0776 0.1889 1.0000 5.000 0.9059 0.00934 0.00342 -0.0763 0.1789 1.0000 5.250 0.9279 0.00949 0.00356 -0.0751 0.1712 1.0000 5.500 0.9492 0.00971 0.00373 -0.0738 0.1609 1.0000 5.750 0.9700 0.00995 0.00391 -0.0724 0.1477 1.0000 6.000 0.9893 0.01029 0.00413 -0.0707 0.1267 1.0000 6.250 1.0044 0.01089 0.00452 -0.0684 0.0897 1.0000 6.500 1.0232 0.01127 0.00484 -0.0667 0.0765 1.0000 6.750 1.0440 0.01154 0.00510 -0.0653 0.0689 1.0000 7.250 1.0759 0.01270 0.00599 -0.0610 0.0252 1.0000 7.500 1.0953 0.01305 0.00634 -0.0594 0.0206 1.0000 7.750 1.1150 0.01337 0.00667 -0.0579 0.0179 1.0000 8.000 1.1342 0.01372 0.00701 -0.0563 0.0158 1.0000 8.250 1.1530 0.01410 0.00738 -0.0547 0.0138 1.0000 8.500 1.1722 0.01441 0.00773 -0.0532 0.0129 1.0000 8.750 1.1897 0.01476 0.00810 -0.0513 0.0120 1.0000 9.000 1.2062 0.01513 0.00848 -0.0492 0.0112 1.0000 9.250 1.2219 0.01557 0.00894 -0.0471 0.0103 1.0000 9.500 1.2386 0.01595 0.00935 -0.0452 0.0098 1.0000 9.750 1.2554 0.01635 0.00979 -0.0433 0.0094 1.0000 10.000 1.2719 0.01677 0.01025 -0.0414 0.0091 1.0000 10.250 1.2880 0.01723 0.01075 -0.0396 0.0087 1.0000 10.500 1.3039 0.01771 0.01125 -0.0378 0.0081 1.0000 10.750 1.3188 0.01827 0.01184 -0.0358 0.0078 1.0000 11.000 1.3323 0.01892 0.01254 -0.0337 0.0073 1.0000 11.250 1.3474 0.01947 0.01314 -0.0319 0.0071 1.0000 11.500 1.3619 0.02006 0.01378 -0.0301 0.0069 1.0000 11.750 1.3755 0.02071 0.01449 -0.0282 0.0067 1.0000 12.000 1.3890 0.02139 0.01523 -0.0264 0.0065 1.0000 12.250 1.4013 0.02214 0.01604 -0.0245 0.0063 1.0000 12.500 1.4132 0.02294 0.01689 -0.0226 0.0061 1.0000 12.750 1.4255 0.02371 0.01771 -0.0209 0.0059 1.0000 13.000 1.4350 0.02471 0.01877 -0.0190 0.0057 1.0000 13.250 1.4449 0.02570 0.01980 -0.0172 0.0054 1.0000 13.500 1.4516 0.02694 0.02113 -0.0152 0.0053 1.0000 13.750 1.4585 0.02822 0.02249 -0.0134 0.0051 1.0000 14.000 1.4657 0.02950 0.02385 -0.0117 0.0050 1.0000 14.250 1.4726 0.03086 0.02529 -0.0102 0.0050 1.0000 14.500 1.4772 0.03246 0.02698 -0.0087 0.0049 1.0000 14.750 1.4831 0.03400 0.02860 -0.0074 0.0048 1.0000 15.000 1.4858 0.03588 0.03058 -0.0061 0.0047 1.0000 15.250 1.4895 0.03777 0.03256 -0.0051 0.0046 1.0000 15.500 1.4892 0.04013 0.03502 -0.0042 0.0046 1.0000 15.750 1.4896 0.04254 0.03753 -0.0036 0.0045 1.0000 16.000 1.4916 0.04492 0.03999 -0.0032 0.0043 1.0000 16.250 1.4895 0.04787 0.04306 -0.0031 0.0043 1.0000 16.500 1.4819 0.05166 0.04697 -0.0033 0.0043 1.0000 16.750 1.4819 0.05461 0.05001 -0.0036 0.0042 1.0000 17.000 1.4703 0.05921 0.05475 -0.0045 0.0042 1.0000 17.250 1.4674 0.06274 0.05835 -0.0053 0.0041 1.0000 17.500 1.4535 0.06802 0.06377 -0.0068 0.0041 1.0000 17.750 1.4431 0.07298 0.06885 -0.0084 0.0041 1.0000 18.000 1.4254 0.07927 0.07528 -0.0107 0.0041 1.0000 18.250 1.4118 0.08514 0.08126 -0.0130 0.0040 1.0000 18.500 1.3881 0.09291 0.08919 -0.0163 0.0040 1.0000 18.750 1.3656 0.10052 0.09694 -0.0196 0.0041 1.0000