XFOIL Version 6.94 Calculated polar for: NACA 16-018 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 ------- -------- --------- --------- -------- ------- ------- 0.500 0.0938 0.02602 0.02084 -0.0066 0.8521 0.8878 1.000 0.1292 0.02616 0.02097 -0.0026 0.8397 0.8962 1.500 0.1750 0.02637 0.02120 -0.0013 0.8312 0.8993 2.000 0.2337 0.02477 0.01962 -0.0030 0.8237 0.9020 2.500 0.3212 0.02284 0.01772 -0.0101 0.8140 0.9035 3.000 0.4040 0.02077 0.01566 -0.0157 0.7913 0.9048 3.500 0.4386 0.02018 0.01512 -0.0120 0.7605 0.9078 4.000 0.5442 0.01945 0.01287 -0.0225 0.4741 0.9072 4.500 0.5188 0.02099 0.01344 -0.0067 0.2772 0.9119 5.000 0.5029 0.02330 0.01462 0.0059 0.0810 0.9176 5.500 0.5251 0.02450 0.01572 0.0114 0.0493 0.9238 6.000 0.5492 0.02545 0.01681 0.0165 0.0483 0.9335 6.500 0.5544 0.02801 0.01952 0.0259 0.0480 0.9489 7.000 0.5999 0.02817 0.01984 0.0260 0.0478 0.9559 7.500 0.6620 0.02945 0.02129 0.0230 0.0480 0.9588 8.000 0.7324 0.03076 0.02272 0.0186 0.0476 0.9613 8.500 0.7912 0.03155 0.02357 0.0157 0.0445 0.9652 9.000 0.8248 0.03285 0.02517 0.0166 0.0266 0.9709 9.500 0.8556 0.03338 0.02567 0.0182 0.0233 0.9770 10.000 0.8955 0.03382 0.02607 0.0187 0.0202 0.9817 10.500 0.9318 0.03486 0.02713 0.0202 0.0179 0.9851 11.000 0.9625 0.03616 0.02848 0.0226 0.0164 0.9875 11.500 0.9901 0.03728 0.02976 0.0255 0.0161 0.9890 12.000 1.0074 0.03951 0.03216 0.0301 0.0152 0.9897 12.500 1.0166 0.04253 0.03548 0.0358 0.0147 0.9899 13.000 1.0168 0.04624 0.03957 0.0425 0.0146 0.9898 13.500 1.0023 0.05084 0.04461 0.0503 0.0148 0.9898 14.000 0.9669 0.05740 0.05173 0.0589 0.0153 0.9902 14.500 0.9003 0.06848 0.06360 0.0666 0.0161 0.9914 15.000 0.8251 0.08422 0.07997 0.0658 0.0167 0.9946 15.500 0.7519 0.10428 0.10050 0.0568 0.0169 0.9969