XFOIL Version 6.94 Calculated polar for: GOE 143 (MVA H.20) AIRFOIL 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 0.4066 0.01081 0.00355 -0.0773 0.7600 0.0734 0.500 0.4606 0.01049 0.00310 -0.0763 0.7118 0.0821 1.000 0.5141 0.01035 0.00275 -0.0754 0.6465 0.0932 1.500 0.5673 0.01034 0.00250 -0.0746 0.5766 0.1226 2.000 0.6161 0.00861 0.00256 -0.0732 0.5249 1.0000 2.500 0.6702 0.00911 0.00271 -0.0728 0.4814 1.0000 3.000 0.7234 0.00965 0.00293 -0.0722 0.4278 1.0000 3.500 0.7764 0.01022 0.00316 -0.0717 0.3663 1.0000 4.000 0.8282 0.01097 0.00348 -0.0712 0.2841 1.0000 4.500 0.8750 0.01265 0.00424 -0.0704 0.1360 1.0000 5.000 0.9266 0.01351 0.00493 -0.0698 0.1046 1.0000 5.500 0.9741 0.01500 0.00593 -0.0689 0.0292 1.0000 6.000 1.0242 0.01602 0.00691 -0.0679 0.0086 1.0000 6.500 1.0751 0.01681 0.00785 -0.0670 0.0071 1.0000 7.000 1.1246 0.01778 0.00904 -0.0659 0.0069 1.0000 7.500 1.1726 0.01893 0.01049 -0.0645 0.0069 1.0000 8.000 1.2179 0.02040 0.01230 -0.0627 0.0071 1.0000 8.500 1.2583 0.02239 0.01474 -0.0604 0.0075 1.0000 9.000 1.2885 0.02521 0.01789 -0.0569 0.0079 1.0000 9.500 1.3076 0.02874 0.02171 -0.0522 0.0084 1.0000 10.000 1.3173 0.03274 0.02597 -0.0465 0.0089 1.0000 10.500 1.3238 0.03753 0.03101 -0.0410 0.0094 1.0000 11.000 1.3375 0.04363 0.03738 -0.0370 0.0101 1.0000 11.500 1.3519 0.04656 0.04061 -0.0338 0.0111 1.0000 12.000 1.3453 0.05509 0.04989 -0.0299 0.0140 1.0000 12.500 1.3310 0.06440 0.05989 -0.0279 0.0197 1.0000 13.000 1.2846 0.07707 0.07318 -0.0315 0.0223 1.0000