XFOIL Version 6.96 Calculated polar for: E184 (8.33%) 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.500 -0.6757 0.08234 0.08052 0.0187 1.0000 0.0186 -8.250 -0.6884 0.07662 0.07484 0.0135 1.0000 0.0185 -8.000 -0.6955 0.07005 0.06822 0.0086 1.0000 0.0185 -7.750 -0.6950 0.06435 0.06244 0.0053 1.0000 0.0192 -7.500 -0.6922 0.05830 0.05625 0.0027 1.0000 0.0200 -7.250 -0.6767 0.05281 0.05045 0.0012 1.0000 0.0228 -7.000 -0.6638 0.04880 0.04608 0.0005 0.9600 0.0231 -6.750 -0.6791 0.03927 0.03594 0.0025 0.9074 0.0241 -6.500 -0.6679 0.03638 0.03288 0.0038 0.8843 0.0251 -6.250 -0.6510 0.03459 0.03092 0.0048 0.8663 0.0263 -6.000 -0.6330 0.03249 0.02856 0.0058 0.8511 0.0283 -5.750 -0.6038 0.03436 0.03008 0.0073 0.8371 0.0330 -5.500 -0.5966 0.02234 0.01705 0.0104 0.8280 0.0208 -5.250 -0.5737 0.02007 0.01440 0.0116 0.8176 0.0208 -5.000 -0.5496 0.01751 0.01145 0.0127 0.8076 0.0197 -4.750 -0.5245 0.01574 0.00942 0.0136 0.7982 0.0194 -4.500 -0.4992 0.01454 0.00803 0.0144 0.7899 0.0197 -4.250 -0.4736 0.01359 0.00693 0.0151 0.7809 0.0202 -4.000 -0.4481 0.01281 0.00603 0.0157 0.7730 0.0208 -3.750 -0.4223 0.01218 0.00528 0.0163 0.7652 0.0215 -3.500 -0.3961 0.01170 0.00472 0.0167 0.7577 0.0225 -3.250 -0.3700 0.01120 0.00411 0.0172 0.7506 0.0230 -3.000 -0.3439 0.01060 0.00337 0.0177 0.7436 0.0249 -2.750 -0.3172 0.01022 0.00293 0.0182 0.7369 0.0296 -2.500 -0.2913 0.00965 0.00254 0.0186 0.7304 0.0750 -2.250 -0.2662 0.00904 0.00228 0.0189 0.7239 0.1683 -2.000 -0.2414 0.00846 0.00208 0.0193 0.7179 0.2799 -1.750 -0.2169 0.00787 0.00193 0.0197 0.7114 0.4011 -1.500 -0.1951 0.00716 0.00180 0.0207 0.7061 0.5652 -1.250 -0.1747 0.00647 0.00171 0.0224 0.6997 0.7215 -1.000 -0.1294 0.00596 0.00184 0.0196 0.6942 0.9265 -0.750 -0.0755 0.00619 0.00200 0.0144 0.6883 0.9680 -0.500 -0.0332 0.00633 0.00205 0.0115 0.6821 0.9815 -0.250 0.0140 0.00642 0.00203 0.0074 0.6765 0.9901 0.000 0.0602 0.00643 0.00199 0.0034 0.6699 0.9972 0.500 0.1212 0.00641 0.00186 0.0022 0.6572 1.0000 0.750 0.1467 0.00642 0.00179 0.0027 0.6504 1.0000 1.000 0.1727 0.00640 0.00176 0.0030 0.6423 1.0000 1.250 0.1985 0.00641 0.00172 0.0035 0.6350 1.0000 1.500 0.2246 0.00640 0.00170 0.0038 0.6260 1.0000 1.750 0.2508 0.00642 0.00169 0.0042 0.6181 1.0000 2.000 0.2771 0.00643 0.00168 0.0045 0.6096 1.0000 2.250 0.3035 0.00644 0.00170 0.0049 0.5997 1.0000 2.500 0.3298 0.00647 0.00172 0.0052 0.5903 1.0000 2.750 0.3562 0.00650 0.00174 0.0055 0.5804 1.0000 3.000 0.3828 0.00653 0.00178 0.0058 0.5685 1.0000 3.250 0.4093 0.00657 0.00183 0.0061 0.5553 1.0000 3.500 0.4359 0.00662 0.00190 0.0064 0.5411 1.0000 3.750 0.4624 0.00670 0.00197 0.0067 0.5241 1.0000 4.000 0.4891 0.00679 0.00205 0.0070 0.5000 1.0000 4.250 0.5155 0.00698 0.00213 0.0072 0.4568 1.0000 4.500 0.5417 0.00749 0.00232 0.0071 0.3675 1.0000 4.750 0.5674 0.00839 0.00272 0.0068 0.2536 1.0000 5.000 0.5926 0.00936 0.00321 0.0064 0.1508 1.0000 5.250 0.6175 0.01019 0.00370 0.0063 0.0811 1.0000 5.500 0.6416 0.01148 0.00462 0.0064 0.0163 1.0000 5.750 0.6669 0.01199 0.00523 0.0068 0.0144 1.0000 6.000 0.6916 0.01260 0.00593 0.0072 0.0133 1.0000 6.250 0.7157 0.01332 0.00674 0.0077 0.0125 1.0000 6.500 0.7382 0.01434 0.00784 0.0082 0.0114 1.0000 6.750 0.7581 0.01587 0.00949 0.0091 0.0107 1.0000 7.000 0.7801 0.01678 0.01051 0.0099 0.0105 1.0000 7.250 0.8015 0.01784 0.01169 0.0108 0.0101 1.0000 7.500 0.8219 0.01920 0.01318 0.0119 0.0099 1.0000 7.750 0.8416 0.02085 0.01498 0.0131 0.0097 1.0000 8.000 0.8603 0.02299 0.01733 0.0144 0.0098 1.0000 8.250 0.8764 0.02601 0.02065 0.0158 0.0102 1.0000 8.500 0.8945 0.02849 0.02327 0.0170 0.0112 1.0000 12.500 0.5950 0.12240 0.12020 -0.0059 0.0221 1.0000 12.750 0.5909 0.12763 0.12541 -0.0082 0.0214 1.0000