XFOIL Version 6.94 Calculated polar for: GRUMMAN K-2 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 -0.0902 0.01391 0.00945 -0.0202 0.9946 0.7937 0.500 0.0003 0.01436 0.00990 -0.0263 0.9786 0.8030 1.000 0.0777 0.01446 0.01005 -0.0290 0.9612 0.8133 1.500 0.1593 0.01398 0.00963 -0.0334 0.9451 0.8213 2.000 0.2281 0.01330 0.00906 -0.0343 0.9258 0.8254 2.500 0.2961 0.01244 0.00828 -0.0358 0.8949 0.8318 3.000 0.3975 0.01228 0.00724 -0.0427 0.6428 0.8337 3.500 0.4295 0.01404 0.00758 -0.0385 0.3413 0.8410 4.000 0.4689 0.01508 0.00794 -0.0354 0.1990 0.8455 4.500 0.5234 0.01578 0.00834 -0.0357 0.1474 0.8513 5.000 0.5713 0.01630 0.00876 -0.0341 0.1218 0.8558 5.500 0.6274 0.01692 0.00929 -0.0348 0.0969 0.8610 6.000 0.6753 0.01759 0.00987 -0.0333 0.0750 0.8650 6.500 0.7317 0.01850 0.01076 -0.0339 0.0595 0.8700 7.000 0.7774 0.01937 0.01163 -0.0319 0.0508 0.8740 7.500 0.8317 0.02063 0.01293 -0.0321 0.0443 0.8786 8.000 0.8776 0.02214 0.01449 -0.0304 0.0406 0.8822 8.500 0.9282 0.02396 0.01631 -0.0300 0.0372 0.8862 9.000 0.9778 0.02578 0.01847 -0.0293 0.0351 0.8898 9.500 1.0225 0.02826 0.02113 -0.0278 0.0323 0.8938 10.000 1.0666 0.03121 0.02462 -0.0266 0.0303 0.8976 10.500 1.0987 0.03468 0.02859 -0.0234 0.0299 0.9015 11.000 1.1284 0.03952 0.03378 -0.0213 0.0285 0.9053 11.500 1.1187 0.04604 0.04138 -0.0141 0.0279 0.9090 12.000 1.0877 0.05370 0.04974 -0.0069 0.0280 0.9138 12.500 1.0449 0.06208 0.05863 -0.0029 0.0282 0.9176 13.000 0.9896 0.07531 0.07237 -0.0095 0.0283 0.9214