NASA/AMES A-02 AIRFOIL (ames02-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NASA/AMES A-02 AIRFOIL (ames02-il) Reynolds number: 100,000 Max Cl/Cd: 31.2 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-ames02-il-100000-n5.txt Download as CSV file: xf-ames02-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA/AMES A-02 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.7407 0.07140 0.06621 -0.0019 1.0000 0.0262
-8.000 -0.7455 0.06488 0.05946 -0.0049 1.0000 0.0262
-7.750 -0.7385 0.06148 0.05596 -0.0059 1.0000 0.0269
-7.500 -0.7319 0.05695 0.05120 -0.0071 1.0000 0.0275
-7.250 -0.7242 0.05169 0.04557 -0.0081 1.0000 0.0278
-7.000 -0.7135 0.04632 0.03971 -0.0087 1.0000 0.0277
-6.750 -0.6988 0.04147 0.03432 -0.0088 1.0000 0.0277
-6.500 -0.6804 0.03739 0.02969 -0.0087 1.0000 0.0279
-6.250 -0.6591 0.03390 0.02564 -0.0084 1.0000 0.0283
-6.000 -0.6357 0.03080 0.02197 -0.0081 1.0000 0.0291
-5.750 -0.6109 0.02878 0.01964 -0.0079 1.0000 0.0306
-5.500 -0.5851 0.02706 0.01763 -0.0078 1.0000 0.0326
-5.250 -0.5585 0.02509 0.01529 -0.0074 1.0000 0.0338
-5.000 -0.5320 0.02366 0.01380 -0.0073 1.0000 0.0351
-4.750 -0.5050 0.02227 0.01221 -0.0070 1.0000 0.0372
-4.500 -0.4784 0.02121 0.01115 -0.0069 1.0000 0.0401
-4.250 -0.4517 0.02009 0.00997 -0.0066 1.0000 0.0429
-4.000 -0.4252 0.01909 0.00895 -0.0064 1.0000 0.0461
-3.750 -0.3986 0.01825 0.00812 -0.0062 1.0000 0.0510
-3.500 -0.3720 0.01743 0.00731 -0.0061 1.0000 0.0568
-3.250 -0.3453 0.01668 0.00660 -0.0059 1.0000 0.0654
-3.000 -0.3184 0.01597 0.00594 -0.0058 1.0000 0.0792
-2.750 -0.2918 0.01516 0.00532 -0.0058 1.0000 0.1063
-2.500 -0.2661 0.01410 0.00479 -0.0058 1.0000 0.1943
-2.250 -0.2440 0.01254 0.00445 -0.0055 1.0000 0.4479
-2.000 -0.2230 0.01179 0.00447 -0.0037 1.0000 0.6323
-1.750 -0.2014 0.01148 0.00447 -0.0016 1.0000 0.7343
-1.500 -0.1804 0.01125 0.00446 0.0008 1.0000 0.8103
-1.250 -0.1554 0.01111 0.00447 0.0026 1.0000 0.8833
-1.000 -0.1096 0.01110 0.00447 -0.0002 1.0000 0.9577
-0.750 -0.0585 0.01108 0.00436 -0.0050 1.0000 0.9942
-0.500 -0.0347 0.01104 0.00423 -0.0049 1.0000 1.0000
-0.250 -0.0184 0.01106 0.00418 -0.0033 1.0000 1.0000
0.000 -0.0006 0.01114 0.00420 -0.0020 1.0000 1.0000
0.250 0.0453 0.01118 0.00418 -0.0059 0.9856 1.0000
0.500 0.0940 0.01120 0.00416 -0.0100 0.9625 1.0000
0.750 0.1318 0.01124 0.00417 -0.0118 0.9372 1.0000
1.000 0.1614 0.01131 0.00421 -0.0117 0.9113 1.0000
1.250 0.1866 0.01139 0.00424 -0.0105 0.8842 1.0000
1.500 0.2099 0.01144 0.00427 -0.0089 0.8536 1.0000
1.750 0.2328 0.01149 0.00427 -0.0072 0.8196 1.0000
2.000 0.2554 0.01154 0.00426 -0.0053 0.7787 1.0000
2.250 0.2779 0.01163 0.00423 -0.0034 0.7250 1.0000
2.500 0.3002 0.01182 0.00416 -0.0014 0.6473 1.0000
2.750 0.3224 0.01227 0.00413 0.0004 0.5373 1.0000
3.000 0.3463 0.01300 0.00426 0.0013 0.4167 1.0000
3.250 0.3718 0.01376 0.00453 0.0015 0.3213 1.0000
3.500 0.3981 0.01444 0.00488 0.0016 0.2557 1.0000
3.750 0.4248 0.01505 0.00526 0.0016 0.2097 1.0000
4.000 0.4516 0.01564 0.00568 0.0016 0.1756 1.0000
4.250 0.4783 0.01626 0.00614 0.0016 0.1500 1.0000
4.500 0.5052 0.01685 0.00668 0.0017 0.1296 1.0000
4.750 0.5318 0.01751 0.00724 0.0017 0.1139 1.0000
5.000 0.5582 0.01819 0.00785 0.0018 0.1013 1.0000
5.250 0.5847 0.01889 0.00855 0.0019 0.0906 1.0000
5.500 0.6109 0.01963 0.00928 0.0020 0.0820 1.0000
5.750 0.6369 0.02043 0.01006 0.0021 0.0748 1.0000
6.000 0.6629 0.02125 0.01091 0.0023 0.0685 1.0000
6.250 0.6887 0.02216 0.01191 0.0025 0.0630 1.0000
6.500 0.7143 0.02316 0.01299 0.0027 0.0584 1.0000
6.750 0.7396 0.02418 0.01405 0.0029 0.0546 1.0000
7.000 0.7648 0.02527 0.01524 0.0032 0.0512 1.0000
7.250 0.7896 0.02663 0.01678 0.0035 0.0482 1.0000
7.500 0.8140 0.02772 0.01790 0.0036 0.0457 1.0000
7.750 0.8379 0.02948 0.02001 0.0040 0.0431 1.0000
8.000 0.8613 0.03103 0.02170 0.0043 0.0415 1.0000
8.250 0.8831 0.03318 0.02422 0.0046 0.0397 1.0000
8.500 0.9041 0.03511 0.02644 0.0049 0.0379 1.0000
8.750 0.9244 0.03688 0.02838 0.0052 0.0365 1.0000
9.000 0.9377 0.04085 0.03305 0.0058 0.0354 1.0000
9.250 0.9478 0.04487 0.03760 0.0062 0.0346 1.0000
9.500 0.9545 0.04887 0.04206 0.0065 0.0340 1.0000
9.750 0.9622 0.05199 0.04543 0.0068 0.0334 1.0000
10.000 0.9665 0.05527 0.04892 0.0070 0.0328 1.0000
10.250 0.9439 0.06223 0.05640 0.0060 0.0326 1.0000
10.500 0.9186 0.06829 0.06271 0.0042 0.0327 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NASA/AMES A-02 AIRFOIL (ames02-il)