XFOIL Version 6.96 Calculated polar for: RUTAN WING AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -5.000 -0.4388 0.00975 0.00364 -0.0289 0.8536 0.0517 -4.500 -0.3825 0.00935 0.00325 -0.0290 0.8432 0.0579 -4.000 -0.3257 0.00901 0.00293 -0.0291 0.8338 0.0663 -3.500 -0.2685 0.00870 0.00265 -0.0294 0.8243 0.0766 -3.000 -0.2110 0.00843 0.00241 -0.0297 0.8156 0.0889 -2.500 -0.1533 0.00811 0.00218 -0.0300 0.8065 0.1085 -2.000 -0.0956 0.00781 0.00201 -0.0304 0.7977 0.1424 -1.500 -0.0381 0.00729 0.00182 -0.0308 0.7890 0.2299 -0.500 0.0736 0.00547 0.00148 -0.0315 0.7586 0.6793 0.000 0.1315 0.00539 0.00147 -0.0316 0.7371 0.7249 0.500 0.1897 0.00538 0.00150 -0.0317 0.7187 0.7556 1.000 0.2476 0.00544 0.00153 -0.0318 0.6903 0.7803 2.000 0.3591 0.00613 0.00175 -0.0313 0.5233 0.8250 2.500 0.4063 0.00801 0.00238 -0.0306 0.2119 0.8374 3.000 0.4603 0.00861 0.00270 -0.0304 0.1393 0.8502 3.500 0.5152 0.00906 0.00300 -0.0302 0.1048 0.8632 4.500 0.6243 0.00983 0.00366 -0.0297 0.0745 0.8901 5.000 0.6775 0.01021 0.00405 -0.0291 0.0662 0.9049 5.500 0.7294 0.01064 0.00448 -0.0283 0.0593 0.9215 6.000 0.7786 0.01111 0.00496 -0.0269 0.0534 0.9433 6.500 0.8288 0.01141 0.00532 -0.0257 0.0496 0.9775 7.000 0.8874 0.01201 0.00590 -0.0267 0.0457 1.0000 7.500 0.9410 0.01261 0.00646 -0.0266 0.0425 1.0000 8.000 0.9927 0.01331 0.00718 -0.0262 0.0399 1.0000 8.500 1.0442 0.01395 0.00781 -0.0258 0.0375 1.0000 9.000 1.0929 0.01482 0.00871 -0.0249 0.0354 1.0000 9.500 1.1422 0.01553 0.00945 -0.0242 0.0335 1.0000 10.000 1.1861 0.01671 0.01063 -0.0227 0.0315 1.0000