XFOIL Version 6.94 Calculated polar for: FAUVEL 14% AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.5078 0.04588 0.03518 0.0206 1.0000 0.3214 -2.750 -0.4885 0.04473 0.03363 0.0207 1.0000 0.3334 -2.500 -0.4674 0.04355 0.03224 0.0206 1.0000 0.3528 -2.250 -0.4439 0.04206 0.03088 0.0204 1.0000 0.3818 -2.000 -0.4214 0.04027 0.02950 0.0207 1.0000 0.4316 -1.500 -0.0347 0.04498 0.03377 -0.0329 0.9439 1.0000 -1.250 -0.0198 0.04590 0.03441 -0.0333 0.9364 1.0000 -1.000 -0.0071 0.04691 0.03515 -0.0332 0.9315 1.0000 -0.750 -0.0068 0.04768 0.03571 -0.0310 0.9287 1.0000 -0.500 -0.0046 0.04849 0.03633 -0.0291 0.9260 1.0000 -0.250 -0.0014 0.04936 0.03700 -0.0273 0.9236 1.0000 0.000 -0.0017 0.05018 0.03763 -0.0250 0.9243 1.0000 0.250 -0.0041 0.05097 0.03824 -0.0223 0.9281 1.0000 0.500 -0.0032 0.05190 0.03897 -0.0201 0.9315 1.0000 0.750 -0.0158 0.05219 0.03912 -0.0156 0.9389 1.0000 1.000 -0.0200 0.05286 0.03960 -0.0124 0.9461 1.0000 1.250 -0.0358 0.05295 0.03954 -0.0071 0.9600 1.0000 1.500 -0.0699 0.05204 0.03851 0.0021 0.9853 1.0000 1.750 -0.0895 0.05157 0.03788 0.0089 1.0000 1.0000 2.000 -0.0806 0.05251 0.03862 0.0097 1.0000 1.0000 2.250 -0.0714 0.05349 0.03943 0.0106 1.0000 1.0000 2.500 -0.0620 0.05452 0.04028 0.0113 1.0000 1.0000 2.750 -0.0523 0.05559 0.04120 0.0120 1.0000 1.0000 3.000 -0.0424 0.05671 0.04215 0.0126 1.0000 1.0000 3.250 -0.0323 0.05787 0.04317 0.0132 1.0000 1.0000 3.500 -0.0221 0.05908 0.04425 0.0137 1.0000 1.0000 3.750 -0.0116 0.06033 0.04538 0.0141 1.0000 1.0000 4.000 -0.0011 0.06164 0.04656 0.0145 1.0000 1.0000 4.250 0.0097 0.06299 0.04780 0.0148 1.0000 1.0000 4.500 0.0205 0.06439 0.04910 0.0151 1.0000 1.0000 4.750 0.0314 0.06584 0.05045 0.0153 1.0000 1.0000 5.000 0.0423 0.06734 0.05186 0.0154 1.0000 1.0000