AG14 (ag14-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: AG14 (ag14-il) Reynolds number: 50,000 Max Cl/Cd: 34.82 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ag14-il-50000.txt Download as CSV file: xf-ag14-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: AG14
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
------ -------- --------- --------- -------- -------- --------
-7.750 -0.5471 0.10787 0.10137 0.0153 1.0000 0.1712
-7.500 -0.5334 0.10277 0.09629 0.0164 1.0000 0.1777
-7.250 -0.5383 0.10140 0.09503 0.0107 1.0000 0.1854
-7.000 -0.5228 0.09617 0.08981 0.0138 1.0000 0.1947
-6.500 -0.5081 0.08910 0.08287 0.0099 1.0000 0.2170
-6.250 -0.4982 0.08552 0.07934 0.0091 1.0000 0.2321
-6.000 -0.4882 0.08212 0.07592 0.0079 1.0000 0.2501
-5.750 -0.4791 0.07885 0.07271 0.0040 1.0000 0.2694
-5.500 -0.4661 0.07495 0.06886 0.0065 1.0000 0.2880
-5.250 -0.4549 0.07159 0.06553 0.0047 1.0000 0.3118
-5.000 -0.4427 0.06824 0.06223 0.0045 1.0000 0.3387
-4.750 -0.4304 0.06478 0.05883 0.0064 1.0000 0.3672
-4.500 -0.4182 0.06145 0.05556 0.0091 1.0000 0.3990
-4.000 -0.2706 0.03995 0.03206 -0.0437 1.0000 0.1347
-3.750 -0.2321 0.03477 0.02597 -0.0470 1.0000 0.1159
-3.500 -0.2003 0.03135 0.02190 -0.0482 1.0000 0.1168
-3.250 -0.1719 0.02836 0.01866 -0.0484 1.0000 0.1217
-3.000 -0.1410 0.02561 0.01542 -0.0484 1.0000 0.1243
-2.750 -0.1116 0.02338 0.01278 -0.0481 1.0000 0.1365
-2.500 -0.0832 0.02148 0.01067 -0.0475 1.0000 0.1607
-2.250 -0.0561 0.01968 0.00899 -0.0466 1.0000 0.2086
-2.000 -0.0278 0.01735 0.00733 -0.0459 1.0000 0.3467
-1.750 -0.0101 0.01371 0.00587 -0.0411 1.0000 1.0000
-1.500 0.0182 0.01370 0.00516 -0.0410 1.0000 1.0000
-1.250 0.0450 0.01370 0.00467 -0.0408 1.0000 1.0000
-1.000 0.0712 0.01373 0.00434 -0.0405 1.0000 1.0000
-0.750 0.0971 0.01377 0.00411 -0.0402 1.0000 1.0000
-0.500 0.1227 0.01383 0.00395 -0.0399 1.0000 1.0000
-0.250 0.1481 0.01391 0.00386 -0.0396 1.0000 1.0000
0.000 0.1733 0.01401 0.00380 -0.0393 1.0000 1.0000
0.250 0.1983 0.01414 0.00383 -0.0390 1.0000 1.0000
0.500 0.2230 0.01430 0.00393 -0.0388 1.0000 1.0000
0.750 0.2474 0.01450 0.00410 -0.0387 1.0000 1.0000
1.000 0.2715 0.01474 0.00434 -0.0386 1.0000 1.0000
1.250 0.2952 0.01505 0.00467 -0.0386 1.0000 1.0000
1.500 0.3185 0.01545 0.00511 -0.0388 1.0000 1.0000
1.750 0.3409 0.01598 0.00569 -0.0393 1.0000 1.0000
2.000 0.3620 0.01668 0.00648 -0.0401 1.0000 1.0000
2.250 0.3814 0.01766 0.00756 -0.0415 1.0000 1.0000
2.500 0.4836 0.01826 0.00857 -0.0566 0.9439 1.0000
2.750 0.5603 0.01809 0.00874 -0.0639 0.8852 1.0000
3.000 0.5990 0.01804 0.00889 -0.0631 0.8253 1.0000
3.250 0.6235 0.01821 0.00902 -0.0596 0.7678 1.0000
3.500 0.6442 0.01858 0.00927 -0.0557 0.7112 1.0000
3.750 0.6647 0.01909 0.00962 -0.0523 0.6551 1.0000
4.000 0.6858 0.01973 0.01008 -0.0493 0.5996 1.0000
4.250 0.7077 0.02050 0.01075 -0.0469 0.5447 1.0000
4.500 0.7301 0.02139 0.01146 -0.0448 0.4907 1.0000
4.750 0.7528 0.02241 0.01234 -0.0430 0.4384 1.0000
5.000 0.7758 0.02354 0.01332 -0.0413 0.3898 1.0000
5.250 0.7990 0.02481 0.01451 -0.0399 0.3443 1.0000
5.500 0.8224 0.02620 0.01582 -0.0387 0.3031 1.0000
5.750 0.8460 0.02780 0.01744 -0.0375 0.2676 1.0000
6.000 0.8698 0.02966 0.01942 -0.0365 0.2361 1.0000
6.250 0.8929 0.03154 0.02138 -0.0354 0.2072 1.0000
6.500 0.9161 0.03411 0.02418 -0.0345 0.1850 1.0000
6.750 0.9385 0.03632 0.02632 -0.0335 0.1623 1.0000
7.000 0.9591 0.03981 0.03039 -0.0325 0.1467 1.0000
7.250 0.9783 0.04357 0.03455 -0.0316 0.1339 1.0000
7.500 0.9948 0.04716 0.03852 -0.0308 0.1207 1.0000
7.750 1.0041 0.05329 0.04553 -0.0306 0.1179 1.0000
8.000 1.0083 0.05996 0.05292 -0.0312 0.1175 1.0000
8.250 1.0072 0.06690 0.06032 -0.0325 0.1181 1.0000
8.500 1.0025 0.07387 0.06759 -0.0345 0.1193 1.0000
8.750 0.9966 0.08065 0.07445 -0.0368 0.1206 1.0000
9.000 0.9912 0.08723 0.08112 -0.0390 0.1216 1.0000
9.250 0.9879 0.09366 0.08772 -0.0409 0.1225 1.0000
9.500 0.8843 0.12445 0.11801 -0.0853 0.3004 1.0000
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Polar data table (+)
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