XFOIL Version 6.96 Calculated polar for: RAE(NPL) 5213 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 = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.500 -0.4342 0.10209 0.09475 -0.0002 1.0000 0.3501 -8.250 -0.4286 0.09914 0.09185 0.0003 1.0000 0.3657 -7.750 -0.4147 0.09326 0.08608 0.0018 1.0000 0.3996 -7.500 -0.4058 0.09031 0.08318 0.0027 1.0000 0.4174 -7.250 -0.3981 0.08742 0.08033 0.0035 1.0000 0.4345 -7.000 -0.3962 0.08484 0.07783 0.0044 1.0000 0.4494 -6.750 -0.3886 0.08137 0.07441 0.0044 1.0000 0.4528 -6.500 -0.5438 0.05657 0.04934 -0.0378 1.0000 0.1740 -6.250 -0.5375 0.05036 0.04239 -0.0407 1.0000 0.1500 -6.000 -0.5247 0.04585 0.03694 -0.0415 1.0000 0.1363 -5.750 -0.5069 0.04214 0.03316 -0.0408 1.0000 0.1334 -5.500 -0.4886 0.03908 0.02972 -0.0402 1.0000 0.1324 -5.250 -0.4685 0.03637 0.02661 -0.0394 1.0000 0.1326 -5.000 -0.4469 0.03387 0.02374 -0.0386 1.0000 0.1326 -4.750 -0.4242 0.03174 0.02123 -0.0376 1.0000 0.1340 -4.500 -0.4013 0.02984 0.01900 -0.0365 1.0000 0.1377 -4.250 -0.3789 0.02796 0.01714 -0.0353 1.0000 0.1422 -4.000 -0.3558 0.02652 0.01559 -0.0338 1.0000 0.1482 -3.750 -0.3338 0.02514 0.01424 -0.0321 1.0000 0.1572 -3.500 -0.3121 0.02405 0.01309 -0.0302 1.0000 0.1689 -3.250 -0.2922 0.02277 0.01203 -0.0285 1.0000 0.1878 -3.000 -0.2717 0.02120 0.01084 -0.0274 1.0000 0.2264 -2.750 -0.2703 0.01936 0.01158 -0.0203 1.0000 0.6400 -2.500 -0.2796 0.02038 0.01274 -0.0086 1.0000 0.7395 -2.250 -0.2846 0.02064 0.01302 0.0009 1.0000 0.7973 -2.000 -0.0777 0.02180 0.01328 -0.0154 1.0000 0.9805 -1.750 -0.0214 0.02112 0.01235 -0.0225 1.0000 0.9964 -1.500 -0.0100 0.02083 0.01201 -0.0218 1.0000 1.0000 -1.250 -0.0147 0.02067 0.01184 -0.0183 1.0000 1.0000 -1.000 -0.0214 0.02050 0.01168 -0.0145 1.0000 1.0000 -0.750 -0.0295 0.02031 0.01149 -0.0107 1.0000 1.0000 -0.500 -0.0389 0.02006 0.01126 -0.0067 1.0000 1.0000 -0.250 -0.0478 0.01977 0.01096 -0.0029 1.0000 1.0000 0.000 -0.0504 0.01953 0.01069 -0.0002 1.0000 1.0000 0.250 -0.0391 0.01950 0.01058 0.0004 1.0000 1.0000 0.500 -0.0201 0.01964 0.01063 -0.0003 1.0000 1.0000 0.750 0.0023 0.01990 0.01081 -0.0014 1.0000 1.0000 1.000 0.0259 0.02026 0.01109 -0.0027 1.0000 1.0000 1.250 0.0498 0.02070 0.01147 -0.0039 1.0000 1.0000 1.500 0.0733 0.02121 0.01193 -0.0051 1.0000 1.0000 1.750 0.0962 0.02179 0.01248 -0.0063 1.0000 1.0000 2.000 0.1185 0.02245 0.01312 -0.0073 1.0000 1.0000 2.250 0.1399 0.02318 0.01385 -0.0082 1.0000 1.0000 2.500 0.1736 0.02428 0.01498 -0.0115 0.9939 1.0000 2.750 0.2602 0.02601 0.01689 -0.0235 0.9558 1.0000 3.000 0.3994 0.02545 0.01676 -0.0386 0.8782 1.0000 3.250 0.4766 0.02426 0.01594 -0.0435 0.8373 1.0000 3.500 0.5413 0.02219 0.01423 -0.0441 0.7880 1.0000 3.750 0.5818 0.01984 0.01202 -0.0390 0.6984 1.0000 4.000 0.5939 0.02007 0.01046 -0.0302 0.4208 1.0000 4.250 0.6136 0.02213 0.01151 -0.0287 0.3331 1.0000 4.500 0.6420 0.02369 0.01271 -0.0286 0.2929 1.0000 4.750 0.6722 0.02517 0.01396 -0.0287 0.2667 1.0000 5.000 0.7029 0.02678 0.01535 -0.0289 0.2472 1.0000 5.250 0.7320 0.02838 0.01696 -0.0289 0.2312 1.0000 5.500 0.7599 0.03013 0.01885 -0.0288 0.2183 1.0000 5.750 0.7866 0.03211 0.02090 -0.0286 0.2071 1.0000 6.000 0.8128 0.03412 0.02295 -0.0283 0.1970 1.0000 6.250 0.8354 0.03644 0.02567 -0.0276 0.1888 1.0000 6.500 0.8597 0.03872 0.02796 -0.0272 0.1805 1.0000 6.750 0.8784 0.04164 0.03132 -0.0263 0.1753 1.0000 7.000 0.8945 0.04461 0.03479 -0.0252 0.1705 1.0000 7.250 0.9105 0.04783 0.03835 -0.0243 0.1673 1.0000 7.500 0.9302 0.05111 0.04165 -0.0239 0.1634 1.0000 7.750 0.9415 0.05526 0.04610 -0.0231 0.1617 1.0000 8.000 0.9505 0.05974 0.05092 -0.0223 0.1621 1.0000 8.250 0.9182 0.06678 0.05900 -0.0215 0.1717 1.0000 8.500 0.9163 0.07244 0.06483 -0.0217 0.1761 1.0000 8.750 0.6805 0.10991 0.10249 -0.0625 0.4068 1.0000