XFOIL Version 6.96 Calculated polar for: GOE 5K AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.250 -0.5394 0.10499 0.09811 -0.0103 1.0000 0.0790 -8.000 -0.5445 0.10310 0.09634 -0.0128 1.0000 0.0811 -7.750 -0.5471 0.10110 0.09444 -0.0176 1.0000 0.0820 -7.500 -0.5375 0.09536 0.08870 -0.0138 1.0000 0.0863 -7.250 -0.5333 0.09201 0.08539 -0.0150 1.0000 0.0916 -7.000 -0.5314 0.08970 0.08312 -0.0211 1.0000 0.0956 -6.750 -0.5253 0.08496 0.07845 -0.0198 1.0000 0.0997 -6.500 -0.5183 0.08157 0.07506 -0.0209 1.0000 0.1076 -6.000 -0.5023 0.07509 0.06848 -0.0282 1.0000 0.1256 -5.750 -0.4930 0.07159 0.06494 -0.0290 1.0000 0.1400 -5.500 -0.4845 0.06736 0.06076 -0.0278 1.0000 0.1556 -5.250 -0.4754 0.06341 0.05684 -0.0262 1.0000 0.1726 -4.250 -0.3695 0.04606 0.03815 -0.0353 1.0000 0.0626 -4.000 -0.3394 0.04283 0.03424 -0.0353 1.0000 0.0441 -3.750 -0.3195 0.03897 0.03018 -0.0349 1.0000 0.0392 -3.500 -0.2951 0.03592 0.02667 -0.0343 1.0000 0.0349 -3.250 -0.2712 0.03307 0.02340 -0.0336 1.0000 0.0331 -3.000 -0.2463 0.03053 0.02037 -0.0327 1.0000 0.0326 -2.750 -0.2207 0.02818 0.01749 -0.0317 1.0000 0.0331 -2.500 -0.1948 0.02609 0.01494 -0.0307 1.0000 0.0335 -2.250 -0.1685 0.02412 0.01254 -0.0295 1.0000 0.0336 -2.000 -0.1415 0.02244 0.01045 -0.0284 1.0000 0.0360 -1.750 -0.1154 0.02095 0.00872 -0.0274 1.0000 0.0409 -1.500 -0.0895 0.01965 0.00708 -0.0265 1.0000 0.0495 -1.250 -0.0637 0.01856 0.00572 -0.0258 1.0000 0.0679 -1.000 -0.0213 0.01448 0.00427 -0.0282 1.0000 1.0000 -0.750 0.0018 0.01451 0.00364 -0.0271 1.0000 1.0000 -0.500 0.0246 0.01456 0.00328 -0.0262 1.0000 1.0000 -0.250 0.0472 0.01463 0.00303 -0.0254 1.0000 1.0000 0.000 0.0698 0.01472 0.00288 -0.0245 1.0000 1.0000 0.250 0.0923 0.01482 0.00278 -0.0237 1.0000 1.0000 0.500 0.1147 0.01494 0.00279 -0.0230 1.0000 1.0000 0.750 0.1371 0.01508 0.00285 -0.0222 1.0000 1.0000 1.000 0.1592 0.01523 0.00300 -0.0214 1.0000 1.0000 1.250 0.1814 0.01540 0.00322 -0.0207 1.0000 1.0000 1.500 0.2033 0.01559 0.00350 -0.0199 1.0000 1.0000 1.750 0.2252 0.01580 0.00385 -0.0191 1.0000 1.0000 2.000 0.2469 0.01604 0.00429 -0.0183 1.0000 1.0000 2.250 0.2685 0.01630 0.00489 -0.0175 1.0000 1.0000 2.500 0.2900 0.01658 0.00549 -0.0168 1.0000 1.0000 2.750 0.3114 0.01691 0.00623 -0.0160 1.0000 1.0000 3.000 0.3325 0.01727 0.00721 -0.0150 1.0000 1.0000 3.250 0.4908 0.02314 0.01032 -0.0343 0.0350 1.0000 3.500 0.5201 0.02517 0.01263 -0.0337 0.0308 1.0000 3.750 0.5520 0.02740 0.01523 -0.0330 0.0284 1.0000 4.000 0.5820 0.03002 0.01824 -0.0321 0.0282 1.0000 4.250 0.6094 0.03279 0.02146 -0.0310 0.0285 1.0000 4.500 0.6343 0.03573 0.02492 -0.0296 0.0288 1.0000 4.750 0.6569 0.03880 0.02850 -0.0280 0.0290 1.0000 5.000 0.6774 0.04211 0.03228 -0.0264 0.0293 1.0000 5.250 0.6957 0.04570 0.03625 -0.0249 0.0300 1.0000 5.500 0.7156 0.04896 0.04020 -0.0230 0.0331 1.0000 5.750 0.7312 0.05296 0.04459 -0.0215 0.0361 1.0000 6.000 0.7468 0.05710 0.04915 -0.0202 0.0420 1.0000 6.250 0.7619 0.06156 0.05403 -0.0192 0.0523 1.0000 8.750 0.6747 0.09917 0.09315 -0.0273 0.1357 1.0000 9.000 0.6784 0.10367 0.09765 -0.0272 0.1298 1.0000