RG 14 9% AIRFOIL (rg149-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: RG 14 9% AIRFOIL (rg149-il) Reynolds number: 50,000 Max Cl/Cd: 34.12 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rg149-il-50000-n5.txt Download as CSV file: xf-rg149-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: RG 14 9% 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.5414 0.08639 0.07931 -0.0346 1.0000 0.0508
-9.000 -0.5475 0.08166 0.07463 -0.0368 1.0000 0.0506
-8.750 -0.5551 0.07754 0.07055 -0.0377 1.0000 0.0503
-8.500 -0.5632 0.07312 0.06611 -0.0387 1.0000 0.0501
-8.250 -0.5696 0.06884 0.06178 -0.0393 1.0000 0.0498
-8.000 -0.5758 0.06445 0.05727 -0.0395 1.0000 0.0496
-7.750 -0.5793 0.06019 0.05284 -0.0393 1.0000 0.0495
-7.500 -0.5806 0.05594 0.04833 -0.0386 1.0000 0.0495
-7.250 -0.5789 0.05181 0.04383 -0.0377 1.0000 0.0496
-7.000 -0.5743 0.04783 0.03940 -0.0365 1.0000 0.0501
-6.750 -0.5662 0.04438 0.03560 -0.0352 1.0000 0.0518
-6.500 -0.5524 0.04234 0.03351 -0.0340 1.0000 0.0547
-6.250 -0.5387 0.03959 0.03040 -0.0327 1.0000 0.0569
-6.000 -0.5232 0.03665 0.02699 -0.0313 1.0000 0.0582
-5.750 -0.5052 0.03398 0.02384 -0.0298 1.0000 0.0599
-5.500 -0.4852 0.03165 0.02101 -0.0284 1.0000 0.0621
-5.250 -0.4655 0.02992 0.01914 -0.0273 1.0000 0.0665
-5.000 -0.4448 0.02848 0.01751 -0.0261 1.0000 0.0719
-4.750 -0.4221 0.02690 0.01554 -0.0247 1.0000 0.0762
-4.500 -0.4010 0.02551 0.01418 -0.0234 1.0000 0.0814
-4.250 -0.3791 0.02440 0.01288 -0.0220 1.0000 0.0904
-4.000 -0.3582 0.02338 0.01187 -0.0208 1.0000 0.1051
-3.750 -0.3371 0.02233 0.01089 -0.0197 1.0000 0.1252
-3.500 -0.3162 0.02130 0.01000 -0.0188 1.0000 0.1634
-3.250 -0.2972 0.02021 0.00930 -0.0177 1.0000 0.2283
-3.000 -0.2797 0.01907 0.00875 -0.0163 1.0000 0.3343
-2.750 -0.2654 0.01794 0.00846 -0.0138 1.0000 0.4931
-2.500 -0.2538 0.01725 0.00875 -0.0087 1.0000 0.7023
-2.250 -0.2279 0.01731 0.00889 -0.0062 1.0000 0.8690
-2.000 -0.1328 0.01749 0.00857 -0.0177 1.0000 0.9801
-1.750 -0.0965 0.01742 0.00820 -0.0202 1.0000 1.0000
-1.500 -0.0904 0.01738 0.00799 -0.0172 1.0000 1.0000
-1.250 -0.0828 0.01740 0.00785 -0.0144 1.0000 1.0000
-1.000 -0.0731 0.01748 0.00776 -0.0121 1.0000 1.0000
-0.750 -0.0570 0.01763 0.00775 -0.0109 0.9984 1.0000
-0.500 -0.0128 0.01794 0.00786 -0.0150 0.9862 1.0000
-0.250 0.0298 0.01821 0.00797 -0.0186 0.9734 1.0000
0.000 0.0713 0.01843 0.00807 -0.0218 0.9601 1.0000
0.250 0.1119 0.01862 0.00817 -0.0248 0.9464 1.0000
0.500 0.1525 0.01878 0.00826 -0.0277 0.9322 1.0000
0.750 0.1929 0.01890 0.00834 -0.0304 0.9179 1.0000
1.000 0.2338 0.01897 0.00841 -0.0330 0.9034 1.0000
1.250 0.2746 0.01901 0.00846 -0.0355 0.8886 1.0000
1.500 0.3135 0.01901 0.00850 -0.0375 0.8730 1.0000
1.750 0.3511 0.01900 0.00853 -0.0391 0.8569 1.0000
2.000 0.3827 0.01903 0.00860 -0.0395 0.8379 1.0000
2.250 0.4152 0.01904 0.00868 -0.0399 0.8188 1.0000
2.500 0.4488 0.01901 0.00871 -0.0404 0.7999 1.0000
2.750 0.4767 0.01907 0.00883 -0.0399 0.7779 1.0000
3.000 0.5071 0.01907 0.00891 -0.0396 0.7565 1.0000
3.250 0.5331 0.01915 0.00903 -0.0386 0.7312 1.0000
3.500 0.5595 0.01917 0.00908 -0.0374 0.7022 1.0000
3.750 0.5850 0.01921 0.00912 -0.0360 0.6700 1.0000
4.000 0.6081 0.01935 0.00926 -0.0343 0.6362 1.0000
4.250 0.6314 0.01957 0.00948 -0.0328 0.6041 1.0000
4.500 0.6543 0.01984 0.00975 -0.0314 0.5718 1.0000
4.750 0.6763 0.02017 0.01009 -0.0298 0.5367 1.0000
5.000 0.6971 0.02055 0.01042 -0.0280 0.4978 1.0000
5.250 0.7165 0.02100 0.01079 -0.0261 0.4545 1.0000
5.500 0.7348 0.02156 0.01124 -0.0242 0.4079 1.0000
5.750 0.7525 0.02224 0.01178 -0.0223 0.3596 1.0000
6.000 0.7697 0.02307 0.01248 -0.0205 0.3124 1.0000
6.250 0.7868 0.02403 0.01328 -0.0189 0.2697 1.0000
6.500 0.8043 0.02510 0.01421 -0.0175 0.2336 1.0000
6.750 0.8223 0.02626 0.01527 -0.0162 0.2044 1.0000
7.000 0.8401 0.02748 0.01640 -0.0149 0.1801 1.0000
7.250 0.8590 0.02878 0.01771 -0.0138 0.1605 1.0000
7.500 0.8773 0.03015 0.01910 -0.0126 0.1433 1.0000
7.750 0.8971 0.03161 0.02064 -0.0116 0.1293 1.0000
8.000 0.9165 0.03308 0.02224 -0.0105 0.1166 1.0000
8.250 0.9361 0.03465 0.02398 -0.0095 0.1058 1.0000
8.500 0.9555 0.03637 0.02579 -0.0085 0.0973 1.0000
8.750 0.9725 0.03795 0.02753 -0.0074 0.0887 1.0000
9.000 0.9910 0.04025 0.03015 -0.0063 0.0821 1.0000
9.250 1.0072 0.04234 0.03246 -0.0052 0.0763 1.0000
9.500 1.0189 0.04477 0.03522 -0.0037 0.0704 1.0000
9.750 1.0288 0.04736 0.03814 -0.0022 0.0656 1.0000
10.000 1.0416 0.04976 0.04050 -0.0012 0.0617 1.0000
10.250 1.0378 0.05325 0.04464 0.0012 0.0588 1.0000
10.500 1.0330 0.05642 0.04819 0.0033 0.0556 1.0000
10.750 1.0269 0.05935 0.05136 0.0053 0.0535 1.0000
11.000 1.0194 0.06247 0.05467 0.0070 0.0522 1.0000
11.250 1.0195 0.06523 0.05741 0.0080 0.0501 1.0000
11.500 1.0004 0.06955 0.06197 0.0084 0.0498 1.0000
11.750 0.9799 0.07449 0.06714 0.0075 0.0498 1.0000
12.000 0.9581 0.08026 0.07309 0.0052 0.0499 1.0000
|
Polar data table (+)
Polar graphs
<< Back to RG 14 9% AIRFOIL (rg149-il)