XFOIL Version 6.94 Calculated polar for: nabab 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.008 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4232 0.06256 0.04663 0.0149 0.9999 0.5272 -2.750 -0.3986 0.05933 0.04333 0.0138 0.9999 0.5442 -2.500 -0.3721 0.05610 0.04015 0.0135 0.9999 0.5704 -2.250 -0.3442 0.05283 0.03712 0.0136 0.9999 0.6095 -2.000 -0.3113 0.04911 0.03411 0.0143 0.9999 0.6799 -1.750 -0.1166 0.04363 0.02580 -0.0227 0.9999 1.0001 -1.500 -0.0867 0.04366 0.02422 -0.0218 0.9999 1.0001 -1.250 -0.0618 0.04366 0.02327 -0.0203 0.9999 1.0001 -1.000 -0.0381 0.04368 0.02261 -0.0189 0.9999 1.0001 -0.750 -0.0149 0.04372 0.02213 -0.0175 0.9999 1.0001 -0.500 0.0081 0.04379 0.02178 -0.0161 0.9999 1.0001 -0.250 0.0310 0.04387 0.02154 -0.0149 0.9999 1.0001 0.000 0.0539 0.04398 0.02141 -0.0137 0.9999 1.0001 0.250 0.0769 0.04412 0.02137 -0.0125 0.9999 1.0001 0.500 0.0999 0.04427 0.02144 -0.0115 0.9999 1.0001 0.750 0.1231 0.04445 0.02161 -0.0105 0.9999 1.0001 1.000 0.1465 0.04466 0.02189 -0.0096 0.9999 1.0001 1.250 0.1699 0.04491 0.02230 -0.0088 0.9999 1.0001 1.500 0.1935 0.04520 0.02285 -0.0081 0.9999 1.0001 1.750 0.2168 0.04558 0.02358 -0.0077 0.9999 1.0001 2.000 0.2392 0.04611 0.02456 -0.0075 0.9999 1.0001 2.250 0.2589 0.04695 0.02593 -0.0076 0.9999 1.0001 2.500 0.2709 0.04850 0.02796 -0.0080 0.9999 1.0001 2.750 0.2676 0.05136 0.03098 -0.0084 0.9999 1.0001 3.000 0.2592 0.05476 0.03419 -0.0088 0.9999 1.0001 3.250 0.2570 0.05773 0.03695 -0.0094 0.9999 1.0001 3.500 0.2594 0.06038 0.03944 -0.0100 0.9999 1.0001 3.750 0.2644 0.06283 0.04177 -0.0106 0.9999 1.0001 4.000 0.2712 0.06517 0.04401 -0.0112 0.9999 1.0001 4.250 0.2793 0.06744 0.04620 -0.0117 0.9999 1.0001 4.500 0.2884 0.06965 0.04835 -0.0122 0.9999 1.0001 4.750 0.2982 0.07182 0.05046 -0.0126 0.9999 1.0001 5.000 0.3087 0.07397 0.05255 -0.0130 0.9999 1.0001