XFOIL Version 6.94 Calculated polar for: AH 79-K-135/20 B 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.2426 0.01572 0.01076 -0.0458 0.7424 0.9355 0.500 0.3002 0.01562 0.01061 -0.0463 0.7348 0.9492 1.000 0.3783 0.01525 0.01015 -0.0507 0.7292 0.9553 1.500 0.4454 0.01476 0.00958 -0.0528 0.7245 0.9628 2.000 0.5237 0.01463 0.00943 -0.0579 0.7189 0.9666 2.500 0.5812 0.01441 0.00927 -0.0592 0.7075 0.9737 3.000 0.6578 0.01380 0.00873 -0.0636 0.7005 0.9773 3.500 0.7282 0.01284 0.00773 -0.0663 0.6938 0.9817 4.000 0.7935 0.01260 0.00761 -0.0690 0.6818 0.9863 4.500 0.8651 0.01144 0.00652 -0.0718 0.6655 0.9911 5.000 0.9320 0.01126 0.00650 -0.0749 0.6511 0.9959 5.500 0.9911 0.01042 0.00528 -0.0750 0.5736 1.0000 6.000 1.0235 0.01083 0.00557 -0.0710 0.5325 1.0000 6.500 0.9850 0.01300 0.00661 -0.0543 0.3482 1.0000 7.000 0.9949 0.01381 0.00734 -0.0469 0.3115 1.0000 7.500 0.9958 0.01604 0.00908 -0.0399 0.2099 1.0000 8.000 0.9641 0.02192 0.01366 -0.0321 0.0196 1.0000 8.500 0.9910 0.02405 0.01605 -0.0298 0.0173 1.0000 9.000 1.0150 0.02637 0.01856 -0.0275 0.0170 1.0000 9.500 1.0248 0.02974 0.02219 -0.0241 0.0169 1.0000 10.000 1.0394 0.03290 0.02552 -0.0211 0.0168 1.0000 10.500 1.0526 0.03624 0.02908 -0.0181 0.0171 1.0000 11.000 1.0755 0.03937 0.03237 -0.0149 0.0173 1.0000 11.500 1.1068 0.04212 0.03530 -0.0129 0.0173 1.0000 12.000 1.1367 0.04452 0.03787 -0.0114 0.0182 1.0000 12.500 1.1547 0.04857 0.04227 -0.0093 0.0157 1.0000 13.000 1.1724 0.05305 0.04713 -0.0072 0.0155 1.0000 13.500 1.1789 0.05983 0.05437 -0.0049 0.0163 1.0000 14.000 1.1675 0.06701 0.06203 -0.0034 0.0161 1.0000 14.500 1.1500 0.07422 0.06963 -0.0033 0.0154 1.0000 15.000 1.1249 0.08321 0.07902 -0.0046 0.0150 1.0000 15.500 1.0880 0.09548 0.09172 -0.0085 0.0144 1.0000