XFOIL Version 6.94 Calculated polar for: NPL AIRFOIL FROM ARC CP 1372 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.1317 0.00741 0.00276 -0.0540 0.9105 0.6359 0.500 0.1940 0.00724 0.00258 -0.0547 0.8862 0.6462 1.000 0.2495 0.00716 0.00245 -0.0539 0.8503 0.6538 1.500 0.3042 0.00724 0.00234 -0.0528 0.7996 0.6618 2.000 0.3534 0.00736 0.00231 -0.0506 0.7207 0.6709 2.500 0.3847 0.00868 0.00248 -0.0453 0.4460 0.6809 3.000 0.4195 0.01045 0.00305 -0.0417 0.1573 0.6917 3.500 0.4657 0.01151 0.00369 -0.0399 0.0752 0.7038 4.000 0.5151 0.01213 0.00431 -0.0384 0.0603 0.7165 4.500 0.5647 0.01278 0.00497 -0.0370 0.0538 0.7308 5.000 0.6130 0.01354 0.00576 -0.0354 0.0494 0.7468 5.500 0.6611 0.01426 0.00658 -0.0337 0.0465 0.7650 6.000 0.7069 0.01539 0.00780 -0.0317 0.0443 0.7859 6.500 0.7550 0.01626 0.00884 -0.0300 0.0421 0.8105 7.000 0.8009 0.01733 0.01004 -0.0280 0.0403 0.8419 7.500 0.8441 0.01890 0.01189 -0.0254 0.0390 0.8920 8.000 0.8990 0.02062 0.01396 -0.0252 0.0377 1.0000 8.500 0.9482 0.02235 0.01579 -0.0243 0.0361 1.0000 9.000 0.9955 0.02532 0.01884 -0.0235 0.0348 1.0000 9.500 1.0318 0.02859 0.02271 -0.0205 0.0341 1.0000 10.000 1.0559 0.03340 0.02823 -0.0162 0.0330 1.0000 10.500 1.0879 0.03537 0.03039 -0.0134 0.0314 1.0000 11.000 1.1161 0.03838 0.03344 -0.0108 0.0302 1.0000 11.500 1.0817 0.04587 0.04191 -0.0009 0.0295 1.0000 12.000 1.0253 0.05315 0.04971 0.0091 0.0295 1.0000 12.500 0.9718 0.06152 0.05844 0.0123 0.0297 1.0000 13.000 0.9188 0.07281 0.06998 0.0082 0.0300 1.0000