NACA 63A010 AIRFOIL (n63010a-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 63A010 AIRFOIL (n63010a-il) Reynolds number: 200,000 Max Cl/Cd: 47.68 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n63010a-il-200000.txt Download as CSV file: xf-n63010a-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 63A010 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.5823 0.08338 0.08013 -0.0137 1.0000 0.0662
-9.750 -0.6032 0.07499 0.07176 -0.0188 1.0000 0.0657
-9.500 -0.6291 0.06725 0.06399 -0.0233 1.0000 0.0642
-7.250 -0.7436 0.03295 0.02651 -0.0113 1.0000 0.0403
-7.000 -0.7268 0.02875 0.02175 -0.0093 1.0000 0.0373
-6.750 -0.7061 0.02587 0.01840 -0.0077 1.0000 0.0363
-6.500 -0.6836 0.02395 0.01627 -0.0066 1.0000 0.0370
-6.250 -0.6609 0.02269 0.01488 -0.0058 1.0000 0.0392
-6.000 -0.6375 0.02119 0.01321 -0.0048 1.0000 0.0404
-5.750 -0.6142 0.01976 0.01166 -0.0037 1.0000 0.0415
-5.500 -0.5916 0.01862 0.01042 -0.0025 1.0000 0.0430
-5.250 -0.5711 0.01731 0.00908 -0.0011 1.0000 0.0449
-5.000 -0.5529 0.01615 0.00797 0.0004 1.0000 0.0495
-4.750 -0.5328 0.01545 0.00723 0.0018 1.0000 0.0550
-4.500 -0.5157 0.01438 0.00621 0.0037 1.0000 0.0630
-4.250 -0.4974 0.01353 0.00539 0.0053 1.0000 0.0816
-4.000 -0.4866 0.01148 0.00440 0.0073 1.0000 0.2550
-3.750 -0.4756 0.00998 0.00413 0.0097 1.0000 0.5131
-3.500 -0.4565 0.00968 0.00413 0.0114 1.0000 0.6016
-3.250 -0.4355 0.00960 0.00416 0.0127 1.0000 0.6519
-3.000 -0.4137 0.00959 0.00421 0.0139 1.0000 0.6862
-2.750 -0.3917 0.00963 0.00429 0.0149 1.0000 0.7175
-2.500 -0.3586 0.00972 0.00439 0.0138 0.9955 0.7513
-2.250 -0.3210 0.00982 0.00457 0.0122 0.9875 0.7838
-2.000 -0.2824 0.00990 0.00466 0.0102 0.9802 0.8069
-1.750 -0.2436 0.00992 0.00465 0.0081 0.9724 0.8246
-1.500 -0.2051 0.00992 0.00463 0.0059 0.9648 0.8387
-1.250 -0.1657 0.00990 0.00461 0.0037 0.9577 0.8507
-1.000 -0.1303 0.00988 0.00458 0.0024 0.9487 0.8626
-0.750 -0.0930 0.00985 0.00455 0.0007 0.9413 0.8741
-0.500 -0.0621 0.00983 0.00453 0.0005 0.9301 0.8862
-0.250 -0.0315 0.00981 0.00451 0.0003 0.9194 0.8984
0.000 0.0000 0.00981 0.00451 0.0000 0.9094 0.9094
0.250 0.0315 0.00981 0.00451 -0.0003 0.8984 0.9194
0.500 0.0621 0.00983 0.00453 -0.0005 0.8862 0.9301
0.750 0.0930 0.00985 0.00455 -0.0007 0.8741 0.9413
1.000 0.1303 0.00988 0.00458 -0.0024 0.8626 0.9487
1.250 0.1656 0.00990 0.00461 -0.0037 0.8507 0.9577
1.500 0.2050 0.00992 0.00463 -0.0059 0.8387 0.9648
1.750 0.2435 0.00992 0.00465 -0.0080 0.8246 0.9724
2.000 0.2823 0.00990 0.00466 -0.0102 0.8069 0.9803
2.250 0.3209 0.00982 0.00457 -0.0122 0.7838 0.9875
2.500 0.3585 0.00971 0.00439 -0.0138 0.7514 0.9955
2.750 0.3914 0.00963 0.00429 -0.0149 0.7175 1.0000
3.000 0.4135 0.00959 0.00421 -0.0138 0.6864 1.0000
3.250 0.4352 0.00960 0.00416 -0.0127 0.6520 1.0000
3.500 0.4562 0.00968 0.00413 -0.0113 0.6019 1.0000
3.750 0.4754 0.00997 0.00413 -0.0096 0.5136 1.0000
4.000 0.4865 0.01147 0.00440 -0.0073 0.2562 1.0000
4.250 0.4972 0.01352 0.00538 -0.0052 0.0817 1.0000
4.500 0.5155 0.01438 0.00621 -0.0036 0.0630 1.0000
4.750 0.5326 0.01545 0.00723 -0.0018 0.0550 1.0000
5.000 0.5527 0.01614 0.00796 -0.0004 0.0495 1.0000
5.250 0.5709 0.01730 0.00907 0.0011 0.0449 1.0000
5.500 0.5915 0.01861 0.01042 0.0026 0.0430 1.0000
5.750 0.6141 0.01976 0.01166 0.0037 0.0415 1.0000
6.000 0.6374 0.02118 0.01321 0.0048 0.0404 1.0000
6.250 0.6609 0.02269 0.01488 0.0058 0.0392 1.0000
6.500 0.6836 0.02396 0.01628 0.0067 0.0370 1.0000
6.750 0.7061 0.02587 0.01840 0.0077 0.0363 1.0000
7.000 0.7268 0.02876 0.02176 0.0093 0.0373 1.0000
7.250 0.7437 0.03296 0.02651 0.0113 0.0403 1.0000
13.500 0.5369 0.14259 0.13919 -0.0055 0.0448 1.0000
13.750 0.5378 0.14760 0.14418 -0.0052 0.0439 1.0000
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Polar data table (+)
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